Types of thrusters:

(1.1) Liquid thrusters.

Liquid propellant systems such as the Space Shuttle Main Engines (SSME) utilize separate liquid oxidizers and fuels. Liquid propellant systems have the capability to be started, throttled (vary thrust), shut-down, and restarted.

Liquid propulsion systems suffer from inherent constraints imposed by fluid dynamics and propulsion system materials limitations providing limited maximum achievable exhaust velocities on the order of 1500 meters/second.

(1.2) Solid thrusters.

Solid propellants have evolved into composite materials made up of metal powders (fuels), oxidizers, catalysts, binders, curing agents, etc. These composite mixtures are processed into various shapes (pellets, star-core-structures, etc . ) and then cured. Solid propellant systems can be started, but are not able to be throttled, shut-down, or restarted.

Solid propellant thrusters have limited maximum achievable exhaust velocities on the order of 6000 meters/second.

(1.3) Solar thermal thrusters.

Solar Thermal Propulsion (STP) systems are capable of superior performance compared to liquid and solid fuel propulsion systems. An STP system exhaust velocity can be as high as 20,000 meters/second, significantly higher than conventional solid or liquid thruster systems. A higher exhaust velocity allows the use of a proportionately lower exhausted mass flux rate for a given thrust value. The lower the exhausted mass flux rate the lower the fuel mass fraction for the accomplishment of a given mission (impulse). STP systems providing long duration and long distance are suitable for interorbital transfer and maneuvering missions.

STP systems consist of two off-axis elliptical solar reflectors focused to an appropriate solar energy absorber apparatus where a gas such as hydrogen is superheated and expelled at high velocity providing thrust. The reflectors are mounted on reflector mounting rings which can rotate 360 degrees, The reflector mounting rings are mounted to a receiver assembly which can rotate 180 degrees. The solar energy absorber is within the receiver assembly. Through the rotation of the reflectors about the receiver assembly, and the rotation of the receiver assembly, solar radiation is continuously focused on the solar energy absorber. In order to specify the components of the STP system, each concern will be addressed individually.

1. Reflector and reflective surface:

Seamless, goveless, accurately-contoured plastic reflectors made of polyimide under development by SRS Technologies of Huntsville, Alabama promise to be suitable for service as reflectors in the STP system. SRS is using polyimide with glass transition temperatures between 220 degrees C and 263 degrees C, compatible with the temperatures encountered in Chemical Vapor Deposition (CVD) coating techniques for aluminum and other metals used in producing the reflective surface.

The reflective surface of the mirror can be applied to the reflector in space through the use of Chemical Vapor Deposition (CVD). Aluminum and alloys of aluminum are deposited through the thermal pyrolysis of aluminum alkyls and other gases which come in contact with the solar-heated reflector surface. Deposition of aluminum at 220 degrees C yields aluminum of 111 crystal structure, grain size near 1 micron,. and >50% reflectance. Improvements in the reflective surface can be achieved through control of deposition rate and conditions, application of optical reflection coatings, deposition of silver onto the aluminum, or other method. A metal reflective surface provides superior reflector performance and rigidizes the inflated mirror surface, alleviating the need for make-up inflatant over time. Special composition materials can be deposited through incorporation of other metal containing species in

the gas stream of the CVD system.

2. Torus and support structure:

The reflector is mounted on the torus and support structure such that the mirror focuses solar radiation into the receiver to the solar energy absorber. An inflatable torus and support structure can be fabricated with kevlar-weave teflon laminate materials. Upon deployment, the torus and support structure would have nickel carbonyl introduce. Solar radiation exposure heats the inflatable, causing pyrolitic deposition of nickel metal on the inside of the inflatable, rigidizing it to produce load-heaving capacity, high-rigidity and high-pointing-accuracy.

3. Gimbaling receiver assembly:

The gimbaling receiver-assembly is made up of the receiver housing, the reflector mounting ring rotation systems, and the rotation system that mates from the receiver housing to the spacecraft. The receiver mechanically points the reflectors to maintain solar energy focus on the solar energy absorber.

4. Solar energy absorber:

The solar energy absorber produces superheated hydrogen with the heat from the absorption of

focused solar energy. Small capillary metal-matrix heat transfer elements may be useful in the construction of solar energy absorbers. To produce these metal-matrix elements entails the following steps: Impregnate and saturate a bundle of organic-fibers with nickel metal, or an alloy to produce a metal-matrix / organic-fibers structure. This structure is heated and the organic fibers are melted or burned out to produce a solar absorber with open micropore tubules. Frit assemblies with capillary micropores fabricated in this manner could be useful in the construction of a suitable solar energy absorber.

5. Pointing and navigation system: In order for the reflectors to remain focused on the solar energy absorber at all times, the navigation and sun sensing and pointing systems must be integrated in real-time. Upon change in attitude to the sun the receiver mechanism will make suitable adjustments to maintain solar radiation pointing accuracy.

(1.4) Electric Arcjet thrusters.

Arcjet thrusters involve the superheating of gases passing through an electrical arc and can be started, throttled, shut-down, and restarted. These thrusters have maximum achievable exhaust velocities on the order of 20,000 meters/second. Arcjet thrusters have recently been chosen for orbital maintenance service on communications satellites.

(1.5) Electric Single Beam Ion thrusters.

Single beam ion propulsion systems accelerate a flux of positive-ions through negatively charged electrical grids and can be started, throttled, shut-down, and restarted. Ion propulsion systems based on single beam ion thruster configurations have to date only been capable of providing satisfactory performance in low thrust missions, generally with exhaust velocities less than 10,000 meters/second. Maximum achievable exhaust velocities are limited by the emission velocities achievable with hollow cathode emitted electron velocities. Exhaust velocity is proportional to the square root of the accelerating voltage, and so the required electrical power levels increase as the square of the exhaust velocity. Ion propulsion has been unable to provide the levels of thrust required to be considered for the role as primary propulsion systems, due to the lack of suitable ultra-high power/mass ratio power supplies. Ion propulsion has been examined both theoretically and experimentally since the 1950's. Past experimental thrusters have been of the metallic ion class, mercury (Hg), or cesium (Cs), or the noble gas variety xenon (Xe). These thrusters have been of the single beam configuration, as opposed to the Dual Beam Ion (DBI) configuration described herein. Single beam thrusters, of the metallic class, exhaust poisonous metal vapors. Single beam thrusters, of the noble gas variety, exhaust rare gases with limited availabilities.

Passive Neutralization Systems:

Passive Neutralization Systems (PNS) encountered on single beam ion propulsion systems are used to neutralize the positively-charged ion beam exhaust through contribution of electrons, producing a neutral-charge exhaust of high velocity vapor. Passive neutralization is achieved through various mechanisms such as: thermionic emission (hollow cathode), or through positive-ion collisions with an electron contributing Cathodic-Neutralizing-Screen (CNS). Thermionic emission neutralization systems maintain spacecraft charge neutrality through the thermal emission of electrons into the positively-charged ion stream to produce neutral ion species. Ultimate Exhaust velocity is limited to the thermionic emission velocity of the electrons. Cathodic-Neutralizing Screen (CNS) neutralizing systems maintain spacecraft charge neutrality through the collision of positive-ions with a negatively charged grid. Exhausted mass flux rates are limited by the constricting neutralizing grid's presence in the exhaust stream. CNS lifetime is limited by damage caused through the collision of positive-ion flux with the neutralizing grid.

(l.6) Electric Dual Beam Ion thrusters.

Ion propulsion systems of the Dual Beam Ion (DBI) configuration, powered by ultra-high power/mass ratio power supplies, can theoretically provide service as primary propulsion systems for vehicles launched from

Earth into space. DBI propulsion systems exhaust environmentally safe water.

Primary propulsion systems are capable of launching spacecraft into orbit from the surface of the Earth.

Ultra-high power/mass ratio power supplies are possible with the advent of superconductivity, and more

specifically the advent of high temperature superconductivity. Superconducting Dual Beam Ion Thrust

systems will be described herein, in comparison to liquid propulsion, solid propulsion, and classical

single beam ion propulsion systems.

Active Neutralization Systems:

Active neutralization distinguishes the Dual Beam Ion thrusters from the single beam ion thruster passive neutralization systems. Active neutralization involves side-by-side oppositely charged ion beams with equivalent exhaust velocities and stoichiometrically balanced charge fluxes. The Dual Beam Ion (DBI)

proton (+) - oxygen (-2) thruster configurations examined produce environmentally safe water as the exhaust. DBI propulsion systems are theoretically able to deliver the large thrust levels required for service as primary propulsion systems.

High power/mass ratio power supplies:

Ion propulsion systems have historically been unable to perform as primary propulsion systems for spacecraft due to the lack of sufficiently high power/mass ratio power supplies with dedicated ion primary propulsion programs of theoretical modeling and experimental development considered unfruitful.

Fortunately, with the advent of superconducting and other technologies theoretical models of machines with remarkable new levels of power/mass ratio are conceivable. Energy sources such as fusion and matter-antimatter annihilation may be realized within the foreseeable future due to the level of technical advancement in the various technologies. Described within are theoretical models and calculations which

indicate that quite favorable results can be achieved through utilization of ultra-high power/mass ratio power supplies and Dual Beam Ion (DBI) thrusters.

Dual Beam Ion (DBI) propulsion systems consist of a positively charged proton (H⁺) beam and a negatively charged oxygen (O^-2) beam, dual accelerator configuration. Dual Beam configuration promotes charge neutrality for the craft and specifically, proton (H⁺) – oxygen (O^-2) beams are chosen due to environmental concern for a non-toxic exhaust. Spacecraft charge neutrality is maintained when beam currents are equivalent, resulting in stoichiometric water (H₂O) exhaust. Recombination of oppositely charged beams, yields water under normal operations. For asteroid mining operations, or for military weapons use, individual beams with different exhaust vectors, which do not collide and recombine can be used as beam mining tools in outer space, or as formidable beam weapons able to dissect all materials.

Accelerators of both varieties are similar in design and operation. Operation is that of linear electrical acceleration of ions, which are axially focused into a coherent pencil beam by use of an axially applied magnetic focusing field. Lengths of the respective accelerator beams are designed such that the beams can be operated in the earth's atmosphere, and if operated correctly even started in atmosphere.

Paschen's rule for dielectric breakdown in gaseous media yields for the earth's atmosphere about 300,000 volts/meter as the minimum expected Breakdown Voltage Separation Product (BVSP). For a proton (H⁺) accelerator voltage of 3*10⁵ volts, an accelerator produced exhaust velocity of 7.57*10⁺⁶ meters/second results. This type of dramatic increase in exhaust velocity is the main advantage of ION PROPULSION - ACTIVE NEUTRALIZATION SWEETENS THE RESULTS BY REMOVING LIMITATIONS OF EXHAUSTED MASS FLUX RATES. Additionally, for an oxygen (O^-2) accelerator voltage of 4.15*10⁶ volts, an accelerator minimum length of 13.8 meters is required.

Atmospheric Startup Sequencing

The desire for use of DBI propulsion systems as primary propulsion orbit achieving systems, leads to the requirement for atmospheric startup, throttling, shutdown and restart capabilities.

Sequencing for atmospheric startup involves the steps:

1. Begin by flushing the individual accelerator tubes with high molecular weight molecules, such as carbon

halides (mostly fluorides), which have markedly higher molecular weights, and so markedly higher dielectric breakdown values when compared to normal atmospheric constituents.

2. Initiate low current axially magnetically focused ion flow in both accelerators. High molecular weight

species provide catastrophic breakdown protection during low current initial flow of ions. Such low current initial ion flow creates an evacuation effect, due to the extremely high 7*10⁶ meters/second, (in the case of the CORRECTION UNDERWAYnewtons thrust specification) exhaust velocity of the ion flow. Under proper operating conditions, the ion flow flushes out the breakdown protection gas and ambient atmosphere. The thrusters once started hinder the re-entry into the tube of ambient atmospheric gas species.

3. Current density for the ion beam may be increased in unison with the axial magnetic focusing field to their maximum values after the high molecular weight dielectric protection gas and any other unwanted initial gaseous constituents have been evacuated by the low current initial ion beam flux. Catastrophic dielectric breakdown phenomena can of course be extremely destructive, and must be avoided by careful design and operation. Startup in atmosphere must be approached with great care and diligence in the correct sequencing of startup events.

Breakdown Voltages for Various Gases

Utilizing the following equations;

Paschen's Law U(br) = p*h = 5.67*(mm Hg)*mm,

U(br) = breakdown minimum product,

p = pressure (mm Hg),

h = separation of electrodes (mm),

alternatively,

U(br) = d*U(br_o) = 0.386*(p/K)*U(br_o),

U(br_o) = minimum breakdown voltage at STP (standard temperature and pressure, 20 degrees celsius, and 760 mm Hg),

d = density (g/cc),

K = temperature (degrees kelvin),

Utilizing Paschen's law, the following examples of

Minimum breakdown voltages are provided;

GAS BREAKDOWN PRODUCT

p * h

air 70%N₂ - l4%O₂ - etc. 326 volts

hydrogen H₂ 195 volts

helium He 19 volts

carbon dioxide C0₂ 293 volts

420 volts

neon Ne 85 volts

argon Ar 137 volts

elegas SF₆ 146 amu 815 volts

freon 12 CCl₂F₂ 120 amu 847 volts

perfluoromethylcyclohexane C₇F₁₄ 350 amu 1956 volts

perfluorodibutylether C₈F₁₈0 54 amu 2445 volts

perfluorodimethylcyclohexane C₈F₁₆ 400 amu 2771 volts

perfluorophenatrene C₁₄F₂₄ 624 amu 3260 volts

Dual Beam Ion propulsion systems calculations:

Calculations of relative performance and specification are presented which are usable in the comparison of STS class craft and DBI class craft. Calculations involving impulse, thrust, energy density, etc., are developed with STS class craft used as the baseline in order to develop understandable performance

comparisons.

Dual Beam Ion propulsion systems developing thrust levels equivalent of STS class craft initial pad lift off thrust are examined. In such modeling the initial pad mass of the DBI class craft is set to be equivalent to the initial pad mass of the STS class craft.

Dual Beam Ion propulsion systems developing thrust levels equivalent to lesser thrust values which are comparable with the various types of thrusters encountered on STS class craft are also set forth.

In order to compare Dual Beam Ion (DBI) craft primary propulsion system performance with that of Space

Shuttle (STS) craft primary propulsion system performance, we shall begin by using the initial lift off thrust level developed by the 3-Space Shuttle Main Engines and 2-Solid Rocket Boosters in unison, 2.84*10⁷newtons, as the initial lift-off thrust level developed by the Dual Beam Ion propulsion system.

In order to examine other thrust levels, comparison with the other types of thrusters encountered on STS class craft, as well as levels of thrust applicable to future large interplanetary craft is provided.

COMPLETE REFORMATTING AND REWRITE UNDERWAY - October 1st, 2010 Completion Date Expected - Rich

DBI Calculations

The following derivation is provided, which is of interest in understanding the DBI propulsion system.

T = (dx/dt)*(dm/dt) = thrust,

t = duration of thrust,

dx/dt = exhaust velocity,

dm/dt = exhausted mass flux rate,

I = T*t = impulse,

I(H⁺) = proton beam current,

I(O^-2) = oxygen Seam current,

I(H⁺) = I(O^-2) criteria for craft neutrality,

U(H⁺) = proton beam acceleration voltage,

U (O^-2) = oxygen beam acceleration,

voltage, U(total) = U(H⁺) + U(O^-2) = total acceleration voltage

voltage,

Classical calculations of acceleration voltages;

U*q = 1/2(m)*(v)²,

U = acceleration voltage,

q = charge,

q(unit) = 1.6021892(46)*10^-19 coulombs,

faraday constant 96,484.56(.27) coulombs/mole electrons,

avogadro's number = 6.022045(31)*10²³ particles/mole,

m = mass of ion,

v = resultant velocity of ion,

Relativistic calculations of accelerations voltages;

U*q = (m₀ )*(c)²* [(((v ²)*(c^-2 ))^-0.5)^-1 ],

m_O = rest mass of ion,

m₀(H⁺) = 1.6737503*10^-27kilograms,

m₀(0^-2) = 2.6568051*10^-26kilograms,

c = speed of light = 2.99792458*10⁸ meters/second = 9.835690893*10⁸ feet/second = 3.540848721*10¹² feet/hour = 6.703613634*10⁸ miles/hour = 1.862114898*10⁵ miles/second,

dm/dt (H⁺) = [I(H⁺)(coulombs/second)] * [1.0446644*10^-8(Kg/coulomb)],

dm/dt (O^-2) = [I(O^-2) (coulombs/second)] * [8.2911706*10^-8(kg./coulomb)],

dm/dt (total) = I*9.335835*10^-8(Kg/coulomb),

1Kilogram is exhausted with the passage of 1.071141467*10⁷(ampere-seconds),

(1.7) Warp Drive thrusters.

Warp-Drive thrusters are likely to involve some form of warping of the space-time around a vessel. This would involve the expansion of space in front of a ship and the contraction of space behind a ship. Unified Field Theory must develop further to be able to theoretically understand how this can be accomplished. In all likelihood, it will involve energy on the order of what you can achieve by matter-antimatter annihilation reactions. In the future, we may be able to "fall" forward and experience "warp-drive" propulsive capabilities. I certainly hope this is true.

In light of the need to at least model the current ultimate in propulsion technology, I would like to suggest a psuedo-warp-drive capability by using a photon push drive. A Gamma Radiation Push type drive will have an exhaust velocity of the speed of light (c), 300,000,000 meters/second. This type of thruster has the ultimate in exhaust velocity and is able to provide the ultimate in performance with the (4.54*¹²) 4,540,000,000,000 watt-hours/kilogram matter+antimatter annihilation energy density.

Matter+antimatter (proton and anti-proton) annihilation produces gamma rays which are focused by a superconductive reflecting contour into a collimated beam, whereby thrust is derived. Reflection can be optimized by utilization of glancing incidence configurations with multiple glancing reflections resulting in minor losses and an organized common exhaust vector beam of gamma rays.

Sources of energy:

(2. 1) Chemical energy.

Chemical reactions as sources of energy are listed below. Energy densities are too low for consideration as energy sources for primary propulsion systems using electric thrusters such as the Dual Beam Ion thrusters described herein.

Chemical Energy Source Energy Density

Watt-hour/Kilogram

1. Ni-Cd batteries 7.7

2. Ag-Zn batteries 36.3

3. H202 decomposition 231.5

4. Gasoline-air 571.5

5. H₂ + Cl₂ = 2HC1 698.8

6. Hydrazine airbreathing systems 1,127

7. 2Li + F₂ = 2LiF 1,343

8. 4P + 0₂ = 2P₂0₅ 2,536

9. 2H₂ + 0₂ = 2H₂0 internal oxidizer 3,660

10. H₂ + F₂ = 2HF 3,735

11. Si + 2F₂=SiF₄ 4,027

12. 2B + 3F₂ = 2BF₃ 4,410

13. 4Al + 3O₂ = 2Al₂O₃ 4,543

14. 2A1 + 3F₂ = 2A1F₃ 4,560

15. 2H = H₂ nacent hydrogen recombination 12,247

16. H₂ airbreathing systems 32,700

17. Nacent H recombination & airbreathing oxidation 45,000

Conventional oxidation is able to provide up to near 3.66*10³ watt-hours/Kg for hydrogen and self-contained oxidizer, 3.27*10⁴ watt-hours/kg for airbreathing hydrogen systems, with the possible future development of nacent hydrogen storage and recombination providing an additional 1.2*10⁴ watt-hours/kg. The storage of nacent hydrogen is done in a magnetic containment vessel. Hybrid utilization of the energy released by the airbreathing oxidation of hydrogen and the energy released by the recombination of nacent hydrogen summed together, produces the highest energy density achievable with conventional systems, near 4.5*10⁴ watt-hours/kg .

(2.2) Solar energy.

Solar radiation is useful in Solar Thermal Propulsion systems and photovoltaic to electricity conversion. Solar radiation intensity at the average distance from the sun to the earth (150,000,000

kilometers) is on the order of 1,400 watts/meter². Relative distances closer and farther away from the sun have higher or lower solar radiation intensities, respectively.

(2.3) Nuclear fission.

Nuclear fission may be useful in space, with continuing investment in the SP-100 nuclear reactor of the DOD, DOE, and NASA. There may be environmental impact risks which will forever make this option just a dream. These energy-densities are included as food for thought.

Energy Source Energy Density

Watt-hour/Kilogram

1. Pu plutonium fission cycle 13,600,000

2. U uranium fission cycle 4,500,000,000

(2.4) Nuclear fusion.

Nuclear fusion is potentially clean, with sufficient energy density to be useful as a source of energy for electric thruster primary propulsion systems.

Fusion Energy Source Energy Density

Watt-hour/Kilogram

1. D + D = 3He + neutron fusion cycle 21,600,000,000

2. Fusion published estimate 22,700,000,000

3. D + D = T + H fusion cycle 26,590,000,000

4. 6D = 2(4He) + 2H + 2neutrons 31,800,000,000

5. D + T = 4He + neutron 93,580,000,000

Fusion reactor research has centered on magnetic confinement and laser confinement techniques in the past. Examination of various experimental results are illustrative of a new potential approach involving inertial confinement.

In research involving fusion of super-heavy nuclei, it has been found that to form a nucleus of 110 amu, two nuclei of 55 amu can be made to collide. The two 55 amu nuclei must possess the correct momentums, in other words , not too hot and not too cold. Collision at the correct momentum allows the two 55 amu nuclei to overcome the coulomb repulsion so that the strong nuclear force can bind them, forming a 110 amu nucleus. The 110 amu nucleus is very unstable and can only be formed through collision of two 55 amu nuclei at the correct momentums for the 110 amu nucleus to be formed, yet be cool enough to be stable, with fissile decay retarded. Light nuclei with less than 26 protons (iron) inherently are more stable than nuclei with more than 26 protons, and so fusion of light elements is expected to be easier than the fusion of unstable nuclei of greater mass.

Examination of typical particle velocities in magnetic confinement plasma fusion systems reveals a bell or gaussian distribution, with many nuclei too cool and a few nuclei too hot. Ideal conditions would involve nuclei with all the correct momentums, where cool nuclei don't quench the fusion reactions, and where hot nuclei don't cause fissile instabilities. Neutral Beam Injection (NBI) has achieved same success as a technique for modifying the velocity distribution in magnetically confined-plasma fusion experiments. Neutral Beam Injection into magnetic confinement systems has been shown to increase the number of hot nuclei, shifting the velocity distributions, producing plasmas where we are closer to fulfillment of the fusion breakeven point (Lawson criteria).

Conventional magnetic confinement plasma fusion systems suffer from a lack of particle energy and motion uniformity, with particles moving in all of the conceivable 4pi steradians of vectors. Colliding beam configurations can provide inertial confinement at a focus with vectoral cohesion of motion of particles.

Potential colliding beam configurations include:

2-beams - head-on collision at a focus (one-dimensional confinement),

3-beams - 120 degrees apart in a plane with collision at a focus (two-dimensional confinement),

4-beams - three-dimensional tetrahedrally oriented collision at a focus (three-dimensional confinement),

6-beams - three pairs of two head-on colliding beams (three-dimensional confinement).

Only the last two options involving 4-beams or 6-beams provide three-dimensional inertial confinement.

The tetrahedral 4-beam orientation requires significantly greater beam currents than the 6-beam

configuration, due to the desirable head-on collision nature of the 6-beam configuration, For purposes of

specifying a system only the 6-beam, lower beam current, configuration will be considered here. Whether

6-neutral beams or 6-positively charged ion beams and 4-negatively charged electron beams are used must depend upon a more in-depth theoretical analysis of the two configurations. (see illustrations of the two types of 6-beam configurations) The collision of beams at a focus, due to the effect of space charge, reduces the effective coulomb repulsion of nuclei to be fused, and so promotes fusion reactions.

In conventional plasma fusion systems of the deuterium + deuterium = helium4 variety, a detrimental

neutron flux is generated. This is due to the unstable nature of a helium4 with 24MeV, which undergoes fissile decay to he!ium3, a neutron, and 3.25MeV of residual energy. In order to control this detrimental neutron flux, the use of neutron-deficient mixed deuterium/hydrogen beam currents will promote the

formation of helium3, and minimize the formation of the unstable helium4 species.

Continuous operation of inertial confinement systems such as the 6-beam configurations, is desirable

to minimize excessive thermal and pressure cycling as is found in pulsed operation systems. The continuous energy production of the inertial confinement systems is compatible with the two, energy to electricity conversion systems, described in sections 3.2 and 3.3.

Poly-well magnetic confinement and electrostatic well confinement (Dr. Robert W. Bussard, EMC²) are also research areas deserving of consideration.

(2.5) Matter-antimatter annihilation.

Matter+antimatter annihilation has the highest energy density of all known energy production systems. The production and storage of antimatter is at the present time very expensive, and difficult to perform,

however technical understanding is improving and it can be expected that technical barriers will be overcame as research continues. Production of antimatter by collision of accelerated beams with metal foils is currently underway, with advanced techniques involving impingement of gamma-rays on certain metal crystals expected to be a practical antimatter production method in the future. The listed energy density is intended to provide an idea of the magnitude of achievable energy densities.

Matter+Antimatter Annihilation Energy Source

Energy Density

Watt-hour/Kilogram

1. Matter+Antimatter annihilation 4,540,000,000,000 (4.54*10¹²)

Matter-antimatter annihilation energy production can either power energy to electricity conversion systems such as described in section. 3.2, or can be used in warp drive propulsion systems as described in section (1.7).

Energy to electricity conversion:

(3.1) Photovoltaics.

Photovoltaics, also known as solar cells, are solid-state devices which convert solar radiation to D.C. electricity. Photovoltaics are made of materials such as silicon, gallium arsenide, and other semiconductors. Photovoltaics have no moving parts and can be expected to have useful lifetimes in space of from 5 years to 50 years, depending on the materials and methods of fabrication.

Photovoltaic radiant solar energy to D.C. power conversion efficiencies are in the range of 10% to 41% (theoretical), depending on the type of device. AM0 (air mass 0) insolation (solar radiation intensity in earth's orbit around the sun) is near 1400 watts/meter², with the average distance to the earth from the sun being 149,476,000 kilometers. Relative distances closer or farther from the sun will have greater or lesser insolation values, and thus power generation densities. Assuming a location near earth's solar orbit distance, and an efficiency of 30%, yields a required collector area of near 2,380 meter²/megawatt generation capacity. Solar energy is incapable of providing suitable energy densities to be considered as a primary, earth orbit achieving, power supply. Importance in outer space operations, is significant, with electric thruster systems requiring DC power.

(3.2) Faraday monopolar generators,

Faraday monopolar generators are chosen to convert shaft torque into electricity in ultra-high power/mass ratio power supply configurations. Faraday monopolar generators have been used as pulse power generators in high-current applications such as thermal-pulse welding, and defense department research programs. The Center for Electromechanics at the University of Austin, Texas has been involved in these programs for many years.

Faraday monopolar generators consist of a disc armature rotor, which spins in an axially oriented magnetic field. For a disc armature rotor spinning clockwise (looking from above) in the horizontal plane and magnetic north oriented upwards, the outer edge of the spinning disc armature rotor will be negative (-),

with the rotational axis positive (+). Faraday monopolar generators operating in continuous-mode

provide D.C. current with no ripple in voltage, making them compatible with superconductors. Superconductors exhibit markedly lower critical current densities and critical magnetic field strengths in the presence of varying electrical and(or) magnetic fields. Conventional D.C. electrical generators exhibit ripple

in the D.C. wave form due to such things as magnetic field non-linearities and discontinuities in current

collection brush assemblies. In order to produce the ultimate in superconducting ultra-high power/mass ratio power supplies, variances in voltage waveform must be eliminated and faraday monopolar generators exhibit the purest D.C. waveforms conceivable from rotating machines. Another reason why Faraday monopolar generators are excellent candidates for service as ultra-high power/mass ratio power generators is due to the ability to build machines with a minimum of stray magnetic flux.

Considerations involved in the design of Faraday monopolar generators are dealt with in the following sections.

(3.2.1) Faraday disc armature rotor:

A Faraday disc armature rotor spins in an axial magnetic field and through the cross product of the

magnetic field and the Faraday disc armature rotor's velocity vector, voltage is generated between the edge

and the axis of the Faraday disc armature rotor.

The Faraday disc armature rotor is a normal conductor and not a superconductor as the Meissner

effect would exclude magnetic field and prevent voltage generation. Joule heating of the Faraday disc armature rotor due to the flow of current will require advanced active cooling methods in order to maintain acceptable Faraday disc armature rotor temperatures. The Faraday disc armature rotor is a metal/graphite-fiber matrix, in order to provide high-conductivity, high-strength, and controllable coefficient of thermal expansion.

(3.2.2) Current Collection Systems:

Conventional brush current collection systems have limited rubbing speeds (listed below) and so are not

useful in Faraday monopolar ultra-high power/mass ratio generators. For a 5 meter diameter Faraday disc

armature rotor spinning at 30,000 rpm, the effective rubbing speed would Se 7,854 meters/second,

significantly greater than the 100 meters/second rubbing speeds possible with conventional brush current

collection systems.

In order to circumvent this restriction, a non-rubbing current collection system where electrons

are emitted from the outer edge of the Faraday disc armature rotor and collected by an anodic electron counter electrode ring is described. With the configuration described above, the outer edge of the Faraday disc armature rotor is negative (-). Emission of electrons from the outer edge of the disc is achieved through a hybridized approach involving:

(3.2.2.1) Laser induced photoemmision from low work-function materials such as:

1. Cesiated (exposed to cesium) surfaces of semiconductors such as Gallium Arsenide.

2. Oxides of the low work function elements, such as those of IIA, IIIB, lanthanides, actinides.

(3.2.2.2) Electric field emission from low work function materials such as those listed in (3.2.2.1).

(3.2.2.3) Centripetal acceleration (inertial) forces.

(3.2.2.4) High-strength magnetic field at the Faraday disc armature rotors outer edge.

High-current operation will require advanced active cooling of the emitting surface at the outer edge of the

Faraday disc armature rotor and advanced active cooling of the anodic (electron collecting) counter electrode ring around the circumference of the Faraday disc armature rotor.

Current collection at the rotation axis (the positive (+) pole) can be accomplished through similar methods or through the use of liquid metal brush assemblies.

Conventional brush current collection systems performance limits:

300 KA/m² @ 40 meters/second solid carbon-copper.

900 KA/m² @ 100 meters/second carbon fiber.

? KA/m² @ 220 to 250 meters/second for liquid brushes.

(3.2.3) Rare earth magnets:

Rare earth magnet discs are placed above and below the rotating Faraday disc armature rotor and provide the initial magnetic field for phase one of the start-up procedure of the Faraday generator. the magnetization

is axial in these rare earth magnets.

High energy product (BH_max, B in gauss, H in oersteds) rare earth magnets are becoming common in

everything from sound speakers, demonstration kits for superconductors, stepping motors, generators, almost any application conventional magnets find use.

The major categories of rare earth magnets are described below:

1. R(M)₅, where R is an element or a mixture thereof, from the rare earths (for example, Pr, Y, Sm, La, Ce, etc.], IIIB, lanthanides.

M is an element or a mixture thereof, from the group of elements known as ferro elements(Fe, Co, Ni, sometimes Mn).

R(M)₅ – SmCo₅ is the most common of the rare earth magnets commercially produced presently.

2. (R)₂(M)₁₄(IIIA)₁, where R is an element or mixture thereof from the lanthanides.

M is an element or a mixture thereof, from the group of elements known as the ferro elements (Fe, Co, Ni, sometimes Mn).

IIIA is in most cases Boron, however, Aluminum is a possible component of a mixture.

(R)₂(M)₁₄(IIIA)₁ - Nd₂Fe₁₄B₁ is the best known example of these magnets, with energy products of up to between 56 and 80 (rumor) million gauss-oersteds.

Research in Japan indicates replacement for Nd by partial fractions of Dy, and Tb results in increased energy products. It can be noted, Dy and Tb have much greater susceptibilities in the elemental state than Nd. This can be explained through the examination of field pinning mechanisms in these materials, seeming to imply the ultimate density of magnetic flux quanta which can be stored is related to the magnetic susceptibility of the rare earth component. Pinning is also assisted by the mixing of elements on a specific site, in other words, a certain fraction of Co and (or) Ni in the Fe vacancy may be added to the ternary in order to achieve higher retained magnetic field intensities.

(3.2.4) Superconducting magnet coils:

Superconducting magnet coils provide the high-strength magnetic fields for full-power (phase 2 power-up) operation of the Faraday monopolar generators. The superconducting coils are made of high-temperature superconducting materials such as yttrium barium copper oxide or bismuth strontium calcium copper oxide. The superconductors are refrigerated with solid-state thermoelectric cryo-sheathing (see section 3.2.7) to below the superconductor's critical temperature during phase-1 of the power-up sequence and during continuous operation of the Faraday monopolar generator.

Hoop stresses during phase-2 of the power-up sequence and during continuous operation can be alleviated through the use of coils made up of segmented multiple axially oriented superconducting sections attached end-to-end with flexible superconducting connections (see section 3.2.5). Segmented coil windings lend themselves to maintenance activities, repair activities and overall machine reliability.

(3.2.5) Superconductive flexible connections:

High-temperature superconductors suffer from brittle fracture limitations and must be connected by superconductive flexible connections in order to alleviate hoop stresses developed during power-up and

during continuous operation of Faraday monopolar generators. Superconductive flexible connections are

composed of dendrites or fibers of high-temperature superconductor embedded in a matrix of conductive polymer and provide means for accommodation of hoop-stresses produced by mechanical, thermal,

electrical and magnetic loads and variations.

Fabrication of flexible superconductive connections is done by embedding longitudinally oriented superconductive dendrites or fibers in a matrix of conductive polymer (such as polyacetylene). For example, the superconductive dendrites or fibers have an aspect ratio of 10:1, with 1 micron diameter and 10 microns length. The dimensional variances of the superconductive coils are tolerated through the relative

movement of the superconductive dendrites or fibers within the flexible polymer conductive matrix.

The refrigeration (and resultant dimensional contraction of the polymer matrix) of the flexible superconductive/polymer matrix to below the critical temperature of the superconductive fibers results in intimate contact along the superconducting fiber's outer longitudinal surfaces, providing conductive paths.

The conductive polymer matrix has a greater coefficient of thermal expansion than the superconductive fibers.

Thermal excursions above the critical temperature of the superconductive fibers cause the expansion of the

conductive polymer matrix, protecting the superconductive magnet coil through discontinuation of intimate contact between the superconductive fibers.

In the normal conducting state the superconductive fibers have poorer bulk conductivity than the conductive polymer matrix, so during cool down but above the critical temperature of the superconductive fibers, the current is shunted through the conductive polymer matrix. Upon reaching, and below the critical temperature of the superconductive fibers, the current is shunted through the superconductive fibers held in intimate contact. Conversely, as a thermal excursion occurs from below to above the critical temperature, the conductive polymer matrix expands, the superconductive fibers discontinue intimate contact, and the current is shunted through the conductive polymer matrix. The ability to shunt current through the conductive polymer matrix reduces the damage caused by joule heating of the flexible superconductive connection above the critical temperature of the superconductive fibers.

(3.2.6) Superconductive magnetic shaft bearings:

Superconductive magnetic shaft bearings are used in the construction of rotating machines such as Faraday

monopolar generators and turbines. Superconductive magnetic shaft bearings allow rotation of a shaft without frictional wear or losses associated with conventional bearing assemblies.

(3.2.7) Thermoelectric cryo-sheathing

Thermoelectric solid-state refrigeration is chosen to cool the superconductors in ultra-high power/mass

ratio power supplies. Thermoelectric cryo-sheathing of superconductors may be an ideal method of providing reliable service, supplementing conventional liquid and gaseous refrigeration systems.

Thermoelectric solid-state devices are made up of multiple stage alternating p-type/n-type semiconductor

junctions . With current flow in one direction heating occurs and with current flow in the opposite direction

cooling occurs.

Thermoelectric devices have been able to provide cooling to 140 degrees Kelvin, and with further research, temperatures below the critical temperatures of high temperature superconductors is expected.

(3.2.8) Housing, shielding and internal supports:

The housing assembly of the Faraday Monopolar Generator provides mechanical support for mounting in the vessel or spacecraft, acts as the magnetic field containment shielding, and provides mechanical support

for internal assemblies such as rare earth magnets, superconducting magnet coils, superconducting bearings,

and current collection assemblies.

Magnetic Field Containment is achieved by means of diamagnetic materials such as bismuth (Bi) metal for

rare earth magnetic field containment, and superconducting materials for superconducting coil

high-strength magnetic field containment. During periods of inoperation, and during phase-1 of the start-up sequencing, the diamagnetic bismuth lining of the housing acts to contain the magnetic field of the rare earth magnets. During phase-2 of start-up sequencing and during full-power operation of the Faraday monopolar generator the superconductive lining of the housing supplements the diamagnetic bismuth layer, containing the combined rare earth magnet and superconductive coil magnetic fields. The superconductive lining of the housing is cooled to below the critical temperature during phase-1 of the start-up sequencing, and is maintained at temperatures below the critical temperature during full power operations.

Toroidal-contour housing curvature provides optimal magnetic flux-line/housinq-surface incidence, such that magnetic field containment is optimized. Superconductors, such as the superconductive lining of

the housing, exhibit two critical magnetic fields, Hc2 the critical magnetic field intensity for normal (into

the surface) flux-line/surface incidence, and Hc3. the critical magnetic field intensity for glancing (parallel

to surface) flux-line/surface incidence.

Flux-line/surface glancing incidence critical field intensity Hc3 (Hc3=Hc2*1.69) is near 1.69 times the

flux-line/surface normal incidence critical field intensity, Hc2.

(3.2.9) Start-up sequencing and operational description:

Faraday monopolar generator start-up sequencing involves two phases. Phase-1 begins with rotation of

the Faraday disc armature rotor in the permanent magnetic field of the rare earth magnets, generating sufficient electricity to power the thermoelectric cryo-sheathing which cools the superconductive electromagnet coils and the superconductive magnetic field containment lining of the housing. When the superconductive materials are cooled to below the critical temperature, phase-2 begins with the power-up of the superconductive electromagnet coils, and the resultant increase in output voltage to full-power levels.

Full-power operation involves the rotation of the Faraday disc armature rotor(s) in the cumulative magnetic field of the rare earth magnets and the superconductive magnet coils, with the resultant current flow through the current collection assemblies, and the full power output of ultra-high power levels of D.C. current and voltage.

(3.3) Fusion plasma direct electrical conversion:

Fusion Reactor Direct Conversion (FRDC) technology capable of harnessing the energy produced in fusion

plasmas involves the utilization of magnetic separation techniques. Magnetic separation in this system involves deflection of the negatively charged (-) electron cloud and the positively charged (+) ionic cloud, such that each impinge on separate electrodes. Deflection is accomplished by slowly decreasing the focusing axial magnetic flux intensity the exhausting fusion plasma experiences on the way cut of the Vacuum Exhaust Port (VEP) of the fusion reactor. Deflection of the exhausted electrons is more severe and the electron cloud is seen to lie in a large diameter ring distribution where the Cathodic Electron Collecting Electrode (CECE) is positioned. Positively charged ions maintain a straighter trajectory, being deflected to a much lesser degree than the ring like electron cloud, because of a positive ions inherently larger inertial mass than an electron. The central positive ion cloud impinges upon the Anodic Ion Neutralization Electrode (AINE), are neutralized, and continue out of the Vacuum Exhaust Port (VEP) into space. The spatially differentiated ring cloud of negatively charged electrons and the central cloud of positively charged ions impinge on their respective electrodes, developing a potential difference which drives current through the load circuitry.

The negatively charged electrons impinge on the Cathodic Electron Collection Electrode (CECE) and are absorbed into a high work function surface, whose static potential is thus driven negative. These electrons then move through the desired load circuit system(s) and emerge at the surface of the Anodic ion Neutralization Electrodes surface, where the negatively charged electrons recombine with and neutralize the positively charged flux of ions which continue on out of the crafts Vacuum Exhaust Port as neutral species. Designation of cathode and anode is done from the viewpoint of craft electrical systems, such that the cathode provides electrons to the load circuitry, analogous to a battery or other source of DC power. To avoid Anodic Ion Neutralizing Electrode corrosion, the AINE surface is covered with a low work function material with preferably; high melting point, low vapor pressure, high ion milling resistance (low ion milling rates), chemical compatibility with the fusion plasma positive ion exhaust stream, and other properties favoring long Anodic Ion Neutralizing Electrode lifetimes.

The Vacuum Exhaust Port (VEP) exhausted mass flux from such a Fusion Reactor Direct Conversion (FRDC) system does not significantly contribute to the thrust of the DBI craft. This is due to significantly lower exhaust velocity and plasma mass flux rate of the fusion reactors VEP exhaust stream, when compared to the much greater exhaust velocity of the Dual Beam Ion Thrusters themselves.

Direct Conversion technology can theoretically achieve up to near 90% thermal to D.C. (direct current) power conversion efficiencies. High power density rotating machines are not required using such Fusion Reactor Direct Conversion (FRDC) technology, with the resultant improvement in simplicity, reliability, and overall power/mass performance ratios.

Systems configurations:

(4.1) Solar Thermal Propulsion system.

Solar Thermal Propulsion (STP) systems involve the apparatus described in section 1.3, powered by solar radiation described in section 2.2. Solar radiation impinges on the reflector surfaces of the STP system and are focused onto the solar absorber within the receiver assembly, causing super-heating of a gas such as hydrogen, with expulsion at high-velocity to produce thrust.

(4.2) Photovoltaics powering Arcjet thrusters.

Photovoltaics, described in section 3.1, generate electricity from solar radiation, described in section 2.2, and power Arcjet thrusters, described in section 1.4.

(4.3) Photovoltaics powering Single Beam ion thrusters.

Photovoltaics, described in section 3.1, generate electricity from solar radiation, described in section 2.2, and power Single Beam Ion thrusters, described in section 1.5.

(4.4) Photovoltaics powering Dual Beam Ion thrusters.

Photovoltaics, described in section 3.1, generate electricity from solar radiation, described in section 2.2, and power Dual Beam Ion thrusters, described in section 1.5.

(4.5) SP-100 nuclear reactor - turbine - generator powering Dual Beam Ion thrusters.

Nuclear fission reactors, described in section 2.3, produce heat which is converted to shaft torque by turbines and converted to electricity by Faraday monopolar generators, described in section 3.2, producing electricity which powers Dual Beam Ion thrusters, described in section 1.6.

(4.6) Fusion reactor - turbine - generator powering Dual Beam Ion thrusters.

Nuclear fusion reactors, described in section 2.4, produce heat which is converted to shaft torque by turbines and converted to electricity by Faraday monopolar generators, described in section 3.2, producing electricity which powers Dual Beam Ion thrusters, described in section 1.6.

(4.7) Fusion reactor - direct electrical conversion powering Dual Beam Ion thrusters.

Nuclear fusion reactors, described in section 2.4, produce heat which is converted to electricity by fusion Reactor Direct Conversion systems, described in section 3.3, producing electricity which powers Dual Beam Ion thrusters, described in section 1.6.

(4.8) Matter-antimatter annihilation - turbine - generator powering Dual Beam Ion thrusters.

Matter-antimatter annihilation, described in section 2.5, produces heat which is converted to shaft torque by turbines and converted to electricity by Faraday monopolar generators, described in section 3.2, producing electricity which powers Dual Beam Ion thrusters, described in section 1.6.

(4.9) Matter-antimatter annihilation Warp-Drive.

Matter-antimatter annihilation, described in section 2.5, produces gamma-rays which are focused into a collimated beam by Warp-Drive thruster systems, described in section 1.7, producing the ultimate in propulsion performance.

Comparative modeling:

(5.1) Space Shuttle -vs- Dual Beam Ion craft.

In order to establish a baseline readily comparable to present space shuttle (STS) class craft, an equivalent pad mass Dual Beam ion (DBI) class craft will be examined. To do this the following STS performance and specifications are provided, and a theoretical model for an equivalent initial pad mass DBI class craft is provided.

STS shuttle class craft performance and specifications:

Space Transportation System (STS), space shuttle class craft exhibit near the following values of performance and specification.

STS class craft are able to attain a maximum orbital height of 1,100km (690miles) (gravitational acceleration .85g = 8.33m/sec²), and putting this in perspective, geosynchronous orbit is 35,900km (22,300miles) above sea level (gravitational acceleration .15g = 1.47m/sec²).

STS class craft are composed of four principle units, an Orbiter, an External Tank (ET), and two Solid Rocket Boosters (SRB). The Orbiter craft carries the crew and payload, and is attached to the large External Tank. The External Tank (ET) contains the liquid oxygen tank in the nose and the liquid hydrogen tank in the rear, used by the three Space Shuttle Main Engines (SSME) in the rear of the Orbiter. The two Solid Rocket Boosters (SRB), are attached to opposing sides of the External Tank.

The orbiter itself has the following dimensions, 37m (121.4ft) long, 17m (55.8ft high), with a wing span of 24m (78.7ft). The fuselage has dimensions of near 5.6m*5.6m*37m (1,190 m³) (1.19*10⁹cc).

The cargo bay in the Orbiter of an STS class craft measures 4.5m (15ft) in diameter, and 18m (60ft) in length, representing a cargo bay volume of 286m³ (2.86*10⁸cc), implying a gravimetric average density of

0.103grams/cc for the cargo bay contents (low-density cargo, or not much cargo).

The External Tank has the following dimensions, 47m (154ft) total length, and 8.4m (27,5ft) in diameter representing an approximate internal volume of 2,600m³ (2.6*10⁹cc).

Solid Rocket Boosters (SRB) have the following dimensions, 45.5m (l49ft) in length, and 3.6m (12ft) in diameter representing internal volumes of 46.3 m³ (4.63*10⁸cc) each.

NOTE: estimating the volumetric density utilizing the above dimensions yields that the 2,046,000 kg of mass inside of the estimated volume of 4,724 m³ (4.724*10⁹cc), results in an average gravimetric craft density of near .43g/cc.

The total impulse (I = T*t, thrust(T), duration(t)), for STS class craft is near 5,300,000,000 newton-seconds.

The STS class craft develops 28,400,000 newtons initial lift off thrust.

Space Shuttle Main Engines (SSME) thrusters together produce 2,500,000,000 newton-seconds impulse, consisting of 4,800,000 newtons thrust for a duration of 510 seconds.

The two Solid Rocket Boosters (SRB) together produce 2,800,000,000 newton-seconds impulse, consisting of 23,400,000 newtons thrust for a duration of 120 seconds.

Propellants for the STS SSME system consist of 102,000 kg (1,500,000 liters) of liquid hydrogen and 616,500 kg (540,000 liters) of liquid oxygen, with a gravimetric mixing ratio of 1:6, or a volumetric mixing ratio of 5:18, respectively.

Propellants for the STS SRB system consist of 176,000 kg Al (aluminum metal powder), 768,130 kg

NH₄C10₄ (ammonium perchlorate), 1,870 kg Fe₂0₃ (ferric oxide catalyst), 154,000 kg binder and curing agent.

From consideration of the exhausted mass and the resultant impulse produced, it can be shown that the

average exhaust velocity of the exhausted fuel is near 2,900 m/sec.

The Orbiter itself has a mass of 75,000 kg.

The combined mass of the Orbiter and the payload carried on board can be up to 104,500 kg.

The empty SRB casings have a mass of 84,100 kg.

The large empty outer SSME fuel tank (ET) has a mass of 35,400 kg.

The maximum payload capacity is 29,500 kg.

Total fueled pre-launch STS pad mass is on the order of 2,045,000 kg.

NOTE: It can be seen that the maximum payload an STS class craft can deliver to an orbital height of 1,100km is only 1.44% of the initial fueled pre-launch STS pad mass.

STS class craft of 2,045,000 kg pad mass, are found to consist of 1,820,000 kg fuel (89% of pad mass) and only 225,000 kg of actual craft and payload (11% of pad mass). This configuration with 89% of pad mass being fuel and only 9.6% of pad mass being craft, with the remaining 1.4% of pad mass being payload, necessitates significant sacrifices in many ways.

Structural integrity of the STS class craft is compromised in order to carry such a large fuel fraction. This compromise between fuel and craft can lead to increased catastrophic failure probability. This questionable craft integrity factor and the associated low payload capacity leave much to be desired. In order to seriously improve spacecraft performance a new type of propulsion system needs to be examined and developed.

Orbital Maneuvering System thrusters:

Two Orbital Maneuvering System (OMS) thrusters are used by the Shuttle Orbiter craft to adjust orbital

trajectories. Each OMS thruster weighs 118kg, is 1.96m (6.43ft) long and can deliver 26,700 newtons of thrust. The exothermic hypergolic (spontaneous reaction upon mixing) reaction of the monomethylhydrazine (fuel) and nitrogen-tetroxide (oxidizer) mixture assures reliable starting of the OMS thrusters. Together the two OMS thrusters can deliver up to 53,400 newtons of thrust for orbital adjustments, which is near 0.51 newtons/kg of orbiter+payload combined craft mass for orbiter

adjustments.

Reaction Control System thrusters:

Reaction Control System (RCS) thrusters are composed of clusters of thrusters in the nose and tail

of the orbiter which serve to control the attitude (yaw,roll,pitch) of the STS class craft. RCS thrusters

of two kinds are encountered, the 38-primary RCS thrusters which are collectively capable of delivering 3,870 newtons of thrust, and the 6-vernier ACS thrusters which are individually capable of delivering 110 newtons of thrust.

Auxiliary Power Units:

Auxiliary Power Units (APU) generate power for the STS class craft hydraulic systems during ascent and descent. Individually, each APU generates near 135 horsepower (100,000 watts), each APU has a mass of 40

kg, resulting in airbreathing APU power generation densities of near 2.5 watts/(gram of APU mass).

Hydrazine is the fuel for these airbreathing turbines, with 134 kg allotted to each of the 3-APUs, resulting in

90 minutes of service for each one. These performance levels represent an extractable energy density of near

1,127 w-h/kg for hydrazine airbreathing oxidation.

STS SSME class fuels:

STS SSME derive thrust by utilization of the energy released during the oxidation of fuels and the resultant

expansion of gases. The energy density obtainable is near 3,660 watt-hours(w-h)/kilogram(kg) of

hydrogen(H₂) and oxygen(O₂). Exhaust velocity is limited as a function of factors such as the inherent energy density of available fuel-oxidizer mixtures, the fluid dynamics of expanding heated gases, the inherent materials limitations of nozzles and other propulsion system components.

In order to achieve greater mission capabilities the impulse (Thrust*duration) must be increased, with a simultaneous decrease in the required fuel fraction of pad mass, and increase in the permitted craft+payload

fraction of pad mass. Thrust (T) is defined as the multiplication product of the exhausted mass flux rate

(dm/dt), and the exhaust velocity (dx/dt = v) of the exhausted mass flux. In order to maintain a constant

level of thrust and achieve a decrease in the required exhausted mass flux, the exhaust velocity produced by

the propulsion system must be increased. Thermal expansion propulsion systems are all limited by fluid

dynamics and materials constraints, thus electrical acceleration is chosen for development as exhaust

velocities can be dramatically increased.

STS class craft performance -vs- equivalent mass DBI class craft.

DBI class craft of near 2,046,000 kg initial pad mass, equivalent to STS class craft initial pad mass,

with only a 10% fuel fraction, as compared to the 89% fuel fraction of the STS class craft, are theoretically

capable of providing near 30 to 200 times the deliverable impulse (5,300,000,000 newton-seconds) of comparable STS class craft.

DBI class craft producing 28,500,000 newtons of thrust, for a duration of !86 seconds, deliver the equivalent impulse of STS class craft (5,300,000,000 newton-seconds).

This is achieved with an exhaust velocity of 20,000,000 meters/second, and an exhausted mass flux rate of near 1.42 kg/second, for the proton(H⁺) - oxygen(O^-2) DBI system.

With 186 seconds duration, it can be shown that the total exhausted mass from the DBI thrusters is of the order of 264.1 kg of H₂0.

Acceleration of this H₂0 is accomplished by use of near 1.37*10¹³ watt-hours of electricity, which for a fusion reactor direct thermal to D.C. power conversion rate of near 10%, yields a total of 6,336 kg of deuterium fuel.

Summing the mass of the exhausted H₂O and the mass of the deuterium, yields on the order of 6,600 kg fuel used by a DBI class craft to produce 5,300,000,000 newton-seconds impulse, which when compared to the initial 200,000 kg fuel fraction (10% initial DBI pad mass), yields on the order of 30.3 times the deliverable impulse of comparable pad mass STS class craft (1.606*10¹¹ newton-seconds).

Similarly, a fusion reactor thermal energy to DC power conversion rate of near 90%, yields a total of 704 kg deuterium used by a DBI class craft to produce the energy for the 5,300,000,000 newton-seconds impulse. Summation of the deuterium (704 kg), and H₂O (264.1 kg), yields 968 kg required fuel fraction, when considered in context of the initial 10% pad mass fuel fraction yields on the order of 206 times the deliverable impulse of comparable pad mass STS class craft (1.092*10¹² newton-seconds).