Interstellar Transportation Exploration
Abstract
An operational concept for developing and implementing interstellar transportation and exploration as well as enhancing travel to and the colonization of the solar planets is described. The Alpha Centauri System can be reached within 10 years after trip initiation from the Solar System. Exploration of the entire Milky Way galaxy is achievable as the system expands from star to star. Vehicle velocity profiles with greater than 0.5 light speed have been tabulated including the power levels required. The method of acquiring this power is outlined. The design uses propellant exhaust rates on the order of a few hundred kilograms (rest mass) per second, expelled at relativistic velocities, and photon sail effects. The assumption is made that this much mass is available in interplanetary and interstellar space or can be made available via soloton-like particle beams radiated from the trip initiating star. High level system requirements and design approaches are included.
OPERATIONAL CONCEPT
This operational concept is for interstellar transportation, solar and extrasolar planet colonization, and the exploration of distant stellar systems. It requires a system comprising 4 (5 if a particle beam is required) independent interfacing major elements as presented in the system description. This concept defines the starting point that is expected to require benefit of the knowledge gained from its evaluation and implementation.
CONSTRAINTS
A. Propellant exhaust velocities (Ve) must be an appreciable fraction (0.3 to 0.5) of the speed of light to minimize the total propellant mass.
B. The propellant exhaust velocities produced by chemical combustion processes are orders of magnitude smaller than 0.3c therefore, chemical processes are ruled out of consideration for interstellar distances.
Fission processes are too mass intensive.
Fusion processes for the required energy levels are currently out of reach and are likely to remain so for 50 or more years.
Matter/antimatter propulsion processes capable of providing energy at the required levels will defy our cleverness for yet another two or more centuries.
Zero point vacuum energy harnessing and wormhole exploitation processes are too heavily invested in imagination to merit serious consideration.
C. The nearly perfectly collimated beamed energy must be sufficient to provide 10^17 to 10^19 watts and the receiving surface must be as small as it can be and accommodate power densities on the order of .01 to 10 megawatts per square meter. The design goal is to limit the mass of the receiving surface including its support structure to less than 99% of the total mass of the vehicle.
D. Sole sail is a soul sale. Plasma rocket propulsion is required to achieve the desired orbit at the destination stellar system.
ASSUMPTIONS
A. Photovoltaic efficiencies of 50% or greater will be available within 20 years.
B. Material and or processes can be selected (or developed) to accommodate the mass to thrust ratios and power density characteristics required by the system descriptions.
C. The system can be extended from each stellar system to the next by either transporting the essential components from “the next star back” or manufacturing them in place where suitable natural platforms (planets and/or their moons) are found or can be assembled.
D. Sufficient material exists in interplanetary and interstellar space to supply the propellant mass required for propulsion. It is the nature of the gravitational forces at play in each stellar system to cause cometary and asteroidal objects and materials to be expelled into galactic orbits thus providing both propellants and hazards.
GENERAL SYSTEM REQUIREMENTS
MAJOR COMPONENTS See link at end of text for notional sketches of major items.
A. POWER BEAM GENERATOR (PBG).
The PBG collects and concentrates solar radiation and transmits energy in a collimated beam to various receiving stations. To avoid solar occultations, each Interstellar Vehicle (IV) PBG will be placed in solar orbit at 0.05AU (4 to 5 million miles from the surface of the sun) in a plane within ten degrees of orthogonal to the line of sight to the destination system.
1. Solar energy collection and distribution subsystem with ~10^19 watt output capability.
2. Frequency converter/beam former for power conversion to manageable frequency bandwidth.
3. Beam collimator and beam direction controller to power beams of appropriate frequency.
4. Plasma chamber receives power from PBG main power to process propellant (solar wind material if feasible) into a plasma state.
5. Magnetic ion separator (mass spectrograph) to direct ions into different paths to the appropriate ion accelerator commensurate with the mass and charge of each via mass/charge discriminators.
6. Ion accelerators (linac type particle accelerator)
receives power from converter and ions from the appropriate mass/charge discriminator and accelerates ions to the desired propellant exhaust velocity.
7. Thrust control system directs plasma propellant to maintain PBG orbit and orientation as well as beam direction.
8. Crew accommodations these stations must be maintained by operators on site to ensure uninterrupted availability of power to IV.
B. EARTH ORBIT FERRY (EOF). The EOF is a multi-ton reusable vehicle with solar energy powered ion drive thruster capable of takeoffs and landings from earth’s surface within the performance envelopes of commercial jet airplanes.
1. Earth geosynchronous orbit PBG similar to 0.05 AU PBG but smaller with less power output.
2. Air intake and compression system to provide material (primarily O2 and N2) to be ionized.
3. Power receiver/converter on vehicle.
4. Plasma chamber receives energy from converter to generate plasma from propellant (atmospheric oxygen and nitrogen etc.,)
5. Magnetic ion separator---directs positive and negative ions into different paths to the appropriate ion accelerator.
6. Ion accelerator (linac type particle accelerator)receives energy from converter and accelerates ions to the desired Ve.
7. Thrust control system interfaces with ion accelerator to control g-loading and vehicle direction. Electron accelerators may suffice for steering propulsion.
8. Beam detection and rider control subsystem interfaces with the thrust control subsystem to maintain power receiving unit centered in the power beam.
C. INTERSTELLAR VEHICLE (IV). The IV contains major subsystems as follows:
1. Items 3 through 8. from B. above--maximize interchangeability with the EOF and PBG (at each level of assembly).
2. Physical shielding (magnetic/mechanical/laser)for protection from colliding objects (proton to 1 kg mass comet/asteroid type objects).
4. Obstacle detection and avoidance system for objects more massive than 1kg. must include auxiliary boost system capable of returning ship to main energy beam. Radar type receiver must be capable of receiving and decoding the time coded signal transmissions originating from the PBG and reflected from the potentially hazardous object to calculate position and velocity of such objects which would collide with the vehicle unless avoidance measures were taken.
5. Crew life support system for recycling O2, H2O,and wastes and for food production .
6. Gravity simulation subsystem for periods of vehicle coasting or for accelerations much less than g. Centrifuges will simulate earthlike gravity.
7. Communication system adapted to Doppler shifts of radiation red and blue shifted by significant fractions of light speed velocities of IV.
8. Propellant material collection system using atmospheric gases scoops (relatively light robotic craft to skim material from tops of atmospheres) for Venus and the gas giant planets and similar secondaries of target systems. Collectors for material from the rings of Saturn, Kuiper belt, and Oort cloud. Collectors for interstellar gas, dust, and Oort-cloud-like-objects, which may be in galactic orbit between the star system of origin and the target, star system (this area requires abnormal cleverness as IV velocity approaches and exceeds 0.1c).
D. EARTH ORBIT FERRY LAUNCH FACILITY. This facility contains major elements as follows:
1. Launch vehicle final assembly and maintenance area.
2. Special runway for 300 meter by 300 meter vehicle (smaller pending creative design solutions to the power density challenge.)
3. Staging and loading area capable of supporting daily launches.
4. Typical personnel and equipment accommodations.
E.PARTICLE BEAM GENERATOR.
Current concepts for the particle beam generator will require a dedicated Power Beam Generator (PBG) located in solar polar orbit as close to the sun as our materials science can accommodate. This PBG will direct a collimated power beam to the most suitable gas giant, probably Jupiter or Saturn, to power a hydrogen mining system. The hydrogen will be collected from near the top of the atmosphere by by photovoltaic powered receivers in balloon suspended units and passed to orbiting beam forming particle accelerators in orbit about the giant above its atmosphere. The particle beam will be directed along the path to be taken by the photon power beam driven ship.
If the giant generates hazardous radiation at the locations most desirable for the various units, robots controlled by humans as near as they can safely be will do the dirty work. Obviously orbiting units synchronous to the giant's rotation would be most desirable.
PRELIMINARY DESIGN
COMPUTATION FORMULAE:
Interstellar vehicle velocity:
(1) Vs = Vo + Ve*ln(Ms/(Ms-q*t)
Where: Vo = initial velocity in meters/sec.
Ve = velocity of propellant in meters per second.
q = propellant exhaust rate in kilograms/second.
Ms = vehicle mass in kilograms.
t = time interval in seconds.
(2) Power required: ke/sec = .5*q*(Ve)^2/(1 sec).
(3) Amperes = (grams/sec)/(number of grams/liter)* (Avagadro's Number) * coulombs/electron charge. (singly ionized atoms/molecules are assumed).
(4) Volts = watts/amperes.
Efficiency of vehicle receiver power conversion is assumed to be 50% of the incident power available at the receivers photovoltaic panel assemblies.
Earth launch vehicle velocity:
(5) Vs = Vo + Ve*ln(Ms/(Ms-q*t) - 9.81*t*sin(CA).
Where CA is the climb angle
Other items are same as for Interstellar Vehicle
Earth gravity simulation centrifuge:
(6) g' = 4*3.14159^2*r/t^2
t= 96.4 seconds per revolution (0.63 rpm) for a 24 meter radius centrifuge for g' = g.
(7) Solar energy available at 0.05 AU: 1300/0.0025
= 520000 watts/square meter at 50% efficiency
= 260000 watts/square meter output from the PBG power receiver.
Sail force is computed assuming that thirteen hundred watts of radiation generate 4.7*10^-6 newtons of force. The sail force generated by 10^18 watts of photon radiation on the IV Receiver (IVR) is 3.615e+9 newtons. From a = F/m = 3.615e+9/6e+9 = 0.6025 m/sec^2 (0.06142 g) acceleration is provided for the 6e+9 kilogram IV from the sail effect of 10^18 watts.
INTERSTELLAR VEHICLE (IV)
The main features of the IV are listed below.
THE MAIN FUSELAGE is a cylinder 100 meters long by 50 meters diameter. This accommodates the power distribution, propellant acquisition and ionization, thrust engines, control cabin, crew support services, and other subsystems of the IV to provide operational and maintenance tasks as necessary.
THE POWER RECEIVER consists of a 4000-meter outside radius and a 100-meter inside radius for smallest ring of the receiver main panels. The separation between the start of concentric rings of panels is also 100 meters thus the width of each panel is 99 meters and allows one meter for the interfacing of successive rings of panel assemblies. Also, sets of 628, 49-meter wide; 125, 39-meter wide; and 12, 8-meter wide panels, totaling 142.972 million M^2, fill in the center of the receiver assembly. The number of panels is computed as pi times the diameter of the inner circumference of each segment times 2 panels per meter along the inner circumference. Except for the central 100 meter section of the receiver, each panel is 99 meters width (radial direction) by 4000 meters length. This configuration provides a total of 3.88e+11 square meters of receiving surface. For the 6e+09 KG IV, the mass allocation is 15.2 grams per square meter of surface including the support structure. The main fuselage will be connected to the Power Receiver by four struts capable of sustaining the acceleration loads and of conducting the 10e+18 or so watts of power from the receiver to the fuselage. The forward ends of the receiver and the fuselage will be fitted with an obstacle collision protection shield. Also, the receiver will include a propellant collection collar mounted on the forward end of the receiver.
THE THRUST ENGINES will be mounted on the side of the fuselage diametrically opposite from the struts connecting the fuselage and the Power Receiver. The interface between the thrust engines and the fuselage will be designed to allow the engines to be rotated with respect to the fuselage. This provides deceleration of the IV to reduce velocity sufficiently to establish orbit at the destination system while keeping the protective shields in place facing the IV direction of motion. Thrust deflection electromagnets will provide the beam centering and minor course correction steering.
PERFORMANCE CHARACTERISTICS
The performance characteristics and thrust requirements for typical "point solutions" are as follows:
INTERSTELLAR VEHICLE
The parameters common to each case are:
IV rest mass = 6000000000 KG.
Receiving surface = 3.88e+11 M^2.
Mass of propellant atom/molecule in atomic mass units = 1 AMU.
The loss of energy due to the red shift of the beam energy as the IV gains velocities approaching 0.8 c is believed to be negligible. Its value is a product of Planck's constant, beam power, and the change in frequency due to the Doppler effect e.g.,
(8) 6.23 * 10e-34* 10e12*0.8*10e17 ~ 10e-5 or ~10 microwatts. [More analysis needed.]
CASE 1
Ve = 0.5 c
Mrel_max = 17 837 064 214 KG
q = 200 KG/SEC
Vs = 9.8 METERS/SEC
As = 0.999 to 0.1372 g
Ds = - 0.0651 to - 0.2012 g
PI = 6.50e+18
Pt = 2.598e+18 WATTS
Pa = 6.7e+06 WATTS/M^2
I = 1.9e+10 AMPERES
V = 1.3e+08 VOLTS
Tpc = 7.67 YEARS
Vs_max = 0.9417 c
CASE 2
Ve = 0.35c
Mrel_max = 9 974 181 517 KG
q = 120 KG/SEC
Vs = 3.33 METERS/SEC
As = 0.339 to 0.2035 g
Ds = - 0.0731 to - 0.1226 g
PI = 1.77e+18
Pt = 7.06e+17 WATTS
Pa = 1.82e+06 WATTS/M^2
I = 1.2e+10 AMPERES
V = 6.1e+07 VOLTS
Tpc = 9.99 YEARS
Vs_max = 0.7988 c
Parameter definition is as follows:
Ve is velocity of thrust particles in meters per second.
q is rate of propellant usage in kilograms per second.
Vs is vehicle velocity gain each second.
As is vehicle acceleration range (thrust + sail) as a fraction of g.
Ds is vehicle deceleration range (- thrust + sail) as a fraction of g.
PI is total power incident on IV Receiver surface
Pt is thrust power in watts.
Pa is power per square meter of receiver surface
I is current required in amperes
V is voltage required in volts.
Tpc is time to reach and orbit Proxima Centauri
Vs_max is maximum IV velocity as a fraction of c.
Mrel_max is relativistic mass at Vs_max
c is the speed of light 300000 km/sec
g is 9.81 meters per second squared.
EARTH ORBIT FERRY
A preliminary design for the Earth Orbit Ferry includes the following characteristics. The fuselage is a box-like structure 300 meters on the sides and 60 meters in height. As viewed from the top, a square area 250 meters on a side and 50 meters deep is allocated to the power receiver which consists of photovoltaic cells configured in panel assemblies 50 meters high by 250 meters long. A total of 500 such panels will be arranged across the 250 meter width such that there are two panels per meter providing a power receiving surface of 6250000 square meters. The entire mass of the EOF is 5000000 kilograms with 4000000 kilograms allocated to the power receiver and its support structure. This allocation results in 0.64 kilograms per square meter of receiver including the support structure. One corner of the EOF forms its 'nose' (apex of the leading edges) and the thrusters will be located near the diagonally opposite corner.
EOF performance characteristics and energy requirements cases:
CASE 1
Ve = 0.0003c
Ms = 5000000 KG
q = 200 KG/SEC
CA = 5 Degrees
Vs = 2.78 METERS/SEC
As = 0.2835 g
PI = 2.03e+12 WATTS
Pt = 8.08E+11 WATTS
Pa = 41891 WATTS/M^2
I = 6.43e+9 AMPERES
V = 6300 VOLTS
Tm = 30 AMU
Tso = 9.1(2.4PF + 6.7 COAST) HRS
Vs_max = 52150 MI/HR
Vd-atm = 108 MILES
Tatm = 0.34 HOURS
Vs-atm = 7452 MILES/HOUR
Vs-sync = 5857MILES/HOUR
Vs-alt = 22287 miles
CASE 2
Ve = 0.0003c
Ms = 5000000 KG
q = 200 KG/SEC
CA = 7 Degrees
Vs = 2.44 METERS/SEC
As = 0.2489 g
PI = 2.03e+12 WATTS
Pt = 8.2e+11 WATTS
Pa = 324000 WATTS/M^2
I = 6.43e+08 AMPERES
V = 6300 VOLTS
Tm = 30 AMU
Tso = 6.95(2.5PF+4.45 COAST) HRS
Vs_max = 48832 MI/HR
Vd-atm = 93 MILES
Tatm = 0.28 HOURS
Vs-atm = 5462 MILES/HOUR
Vs-sync = 5832 MILES/HOUR
Vs-alt = 22245 miles
CA is the climb angle with respect to horizontal.
Tm mass of thrust particles in atomic mass units.
Tso is time to reach synchronous earth orbit.
Vd-atm is a height reference for being outside most of earth's atmosphere.
Tatm is time required to reach Vd-atm.
Vs-atm is vehicle velocity at time Tatm.
Vs-sync is vehicle velocity in approximate synchronous orbit.
Vs-alt is vehicle altitude when in approximate synchronous orbit.
The receiving surface area for each case is 6.25e+06 meter^2.
POWER BEAM GENERATOR (PBG)
The power output requirement of the IV PBG is 1.77e+18/0.9 or 1.97e+18 watts to 6.5e+18/.9 or 7.23e+18 watts for the beamed energy. If this output can be produced from an input 2.4 times as great, 4.73e+18 to 1.74e+19 watts are required. Using 50% efficient 1-volt photovoltaic cell technology and the solar radiation intensity available at 0.05 AU (4650000 miles) of the sun, 520000 watts/meter^2, the receiving surface area of the PBG must be 9.10e+12 to 3.35e+13 meters squared. In accordance with the interchangeability requirement with the Interstellar Vehicle components, 396000 square meter panel subassemblies must be used. Hence the total number of 396000 square meter panels required is 22.98 to 84.60 million. For the low end of the power requirement, a rectangular assembly consisting of 750 of the 396000 m^2 panels with the long dimension of each placed end to end (LDEE), and 30640 short dimension side by side--(SDSS) meets the requirement. The resulting array will be of total dimensions 3000 kilometers by 3034 kilometers. For the high end of the power requirement, a rectangular assembly consisting of 1440 of the 396000 m^2 with LDEE, and 58750 SDSS meets the requirement. The resulting array will be of total dimensions 5760 kilometers by 5817 kilometers. Final arrangement of the panels must accommodate optimal tradeoffs between ease of assembly and volt/ampere distribution.
EOF PBG.
While in quasi-synchronous orbit about the earth, the EOF PBG will intercept solar radiation at the rate of 1300 watts per M^2. Since 2*2.03e+12/0.9 watts are needed to supply the EOF Power Receiver with thrust plus operating power, the EOF PBG must present 2.82*(2*2.03*e+12/0.9) /1300 or 9.8 billion M^2 to solar radiation. If each individual panel is 396000 M^2, 24711 such panels are required to supply the EOF power. The panels can be configured with 25 LDEE by 989 SDSS resulting in an array of total dimensions 100 kilometers by 98 kilometers. Since this will orbit the earth in approximate geo-synchronous orbit, every effort should be made to select its geo-synchronous orbit to isolate it as much as possible from optical telescopes. The final arrangement of the panels must accommodate optimal tradeoffs between ease of assembly and volt/ampere distribution.
EOF PBG ORBIT
The orbit of the quasi-geosynchronous EOF PBG will be synchronous with the earth's rotation but the plane of its orbit will form a small angle with the plane of the earth's equator. The orbit of the EOF PBG will be such that, as viewed from the EOF Launch Facility, it will describe a thin ellipse (one with large eccentricity) centered above the EOF Launch Facility. Its orbit will be at an altitude and with appropriate velocity to maintain the PBG in direct sunlight for as much of the 24 hour day as is feasible (avoid earth's shadow) while remaining in direct line-of-sight with the Launch Facility.
POWER BUDGET
INTERSTELLAR VEHICLE (IV)
The power provided to the 3.882e+11 M^2 of the IV receiving surface varies primarily as a function of q and Ve. For cases 1 and 2 the power allocation is as shown below. This power level provides the acceleration/ deceleration thrust including distribution losses, ionization power, vehicle operating and control power, and crew accommodation power. The rough calculations presented are intended to establish some degree of feasibility. The parameters q, Ve, Ms, Tm, and Tpc each can be adjusted as needed to optimize feasibility. An approximate power allocation for the 1.77e+18 watts is:
50% 0.833e+18 watts loss due to IV solar array inefficiency
8% 0.14e+18 watts loss to IV operating inefficiency
40% 0.71e+18 watts thrust
2% 0.032e+18 watts propellant gathering, ionization, thrust control,obstacle detection, life support,and gravity simulator.
An approximate power allocation for the 6.50e+18 watts is:
50% 3.25e+18 watts loss due to IV solar array inefficiency
8% 0.52e+18 watts loss to IV operating inefficiency
40% 2.60e+18 watts thrust
2% 0.13e+18 watts propellant gathering, ionization, thrust control, obstacle detection, life support, and gravity simulator.
PHOTOVOLTAIC PANEL EFFICIENCY
The angle of the radiated energy in the Power Beam with respect to the receiving surfaces is very small--0.0225 degrees (large angle of incidence). The configuration of the receiving power panels, due to the folds and the multiple mutual reflections from adjacent facing panels, is assumed to perform equal to or better than receiving surfaces perpendicular to the beam. The actual efficiency will be quantified in the preliminary design phase. Also, the power per square meter and current carrying limits of the "one-volt" technology will be monitored for improvements resulting in several megawatts/M^2 capacity while avoiding the destruction of the panels or the reduction of the operating efficiency of their surfaces. The requirement is to lower the receiver mass while increasing the available power. A heat "scavenger" system will be developed to use the heat energy available from the necessary cooling of the receiving panels to improve the net efficiency of total power usage.
IV PBG
Sufficient power to reach Alpha Centauri in 9.99 years requires approximately 4.73e+18 watts to be input to the PBG from the sun. At 36% of this power level, Alpha Centauri can be reached in 14.6 years. At 2.94 times this power level, Alpha Centauri can be reached in 6.37 years. A guess at how the received power of 4.73e+18 watts is to be allocated in terms of units of 10e+18 watts is as follows:
50% 2.37e+18 watts loss to photovoltaic cell inefficiency
8% 0.37e+18 watts loss to operational inefficiency of PBG.
40% 1.90e+18 watts beam output to Interstellar Vehicle
0.5% 0.023e+18 watts spectral purity and collimation maintenance
1.0% 0.047e+18 watts PBG orientation and orbit maintenance
0.5% 0.023e+18 watts crew accommodation and miscellaneous station operation.
EOF PBG
The initial 2.82*(2*2.03*e+12/0.9) = 12.72e+12 watts power usage estimate in units of 10^12 watts is:
50% 6.36e+12 watts loss due to EOF PBG solar array inefficiency
8% 1.02 e+12 watts loss to EOF PBG operating inefficiency
40% 5.08e+12 watts power beamed to the EOF
2% 0.26e+12 watts EOF PBG orientation, beam control, orbit maintenance, and station operation.
The 40% (5.08e+12 watts) beamed to the EOF is allocated as follows:
20% 2.54e+12 watts loss to EOF solar array inefficiency
3% 0.39e+12 watts loss to EOF operating inefficiency
16% 2.03e+12 watts for thrust power
1% 0.12e+12 watts propellant management and vehicle operation
MASS BUDGET
INTERSTELLAR VEHICLE
Total Mass 6e+09 Kilograms.
99% Receiver Energy Acquisition and Support
Structure 5.94e+09 Kilograms (11.88 e+06 tons).
1% Fuselage and Functional Systems 6e+07
Kilograms (1.32e+05 tons).
Allocated as follows:
20% Receiver Protective Shield
1.2e+07 Kilograms (2.64e+04 Tons)
20% Fuselage and Protective Shield
1.2e+07 Kilograms (2.64e+04 Tons)
20% Propellant acquisition and thrust system
1.2e+07 Kilograms (2.64e+04 Tons)
10% Crew, Crew Support and Gravity Simulator
6e+06 Kilograms (1.2e+04 Tons)
20% Tools, Equipment. Propellant Reserves, &
Spares 1.2e+07 KG (2.64e+04 Tons)
10% Shuttle Craft 6e+06 Kilograms (1.2e+04
Tons)
IV PBG
Total Mass = 1.14e+12 KG (computation assumes 520000 watts/M^2, 0.1 KG per square meter of receiving surface including support structure plus 20% of this mass for the remainder of the PBG).
Receiver = 0.1 KG/M^2 * 9.47e+12 M^2 = 9 .47e+11 KG.
Beam Transmitting Antenna 10% of Receiver mass = 0.947e+11 KG
Control Module including shielding from coronal mass ejections
= 10% of Receiver mass = 0.947e+11 KG.
EOF PBG MASS BUDGET
Total Mass 2.5e+09 KG
Receiver 0.1 KG/M^2*20.8 billion M^2 = 2.1e+9 KG
Beam Transmitting Antenna 10% Receiver mass = 0.2e+9 KG
Control Module 10% Receiver mass = 0.2e+9 KG
EARTH ORBIT FERRY
Total Mass = 5e+06 KG
Receiver = 4e+06 KG
Control Module including engines and landing gear = 0.8e+05 KG Payload = 9.2e+05 KG
BENEFITS.
Some of the more desirable potential benefits of the program are:
1. Facilitation of solar system exploration and colonization by improving ease and speed of transportation and by supplying energy to remote solar sites (Mars and beyond) via the collimated power beams included as part of this system description. The terraforming of:
a. VENUS: Use microbial psychrophilic acidophiles (euplotes antarcticus, polaromonas vacuolata, etc.,) genetically modified, if and to the extent required, to break down the high altitude cold sulfuric acid clouds and barophilic hyperthermophiles [e.g., methanopyrus kandleri, pyrolobus fumarii, etc., ] to break down the lower altitude hot CO2 respectively and:
b. MARS: Beam the necessary solar energy and transfer the water, if necessary, from the lesser moons of the gas giant planets or from the Saturnian rings, KBO, etc.,. Transfer 5 bars of CO2 and N2 from Venus.
Solar system ferry times (sail effects included ) to orbits in solar system assumes earth and object of interest (except the Moon) are within 20 degrees of inferior conjunction (add 3 days for worse case (superior conjunction) to reach object). Stopping times are included as is powered flight at indicated acceleration/deceleration(a/d ).
.Target.................... a = 1.02 g...................... a = 0.677g
................................. d = -0.955 g....................d = -0.511
Moon...................... 3.4 Hours ..................... 4.3 Hours
Mercury...................2.3 Days........................2.84 Days
Venus or Mars.......2.1 Days........................2.5 Days
Jupiter.....................6.1 Days .......................15 Days
Saturn .....................9.5 Days ......................22.5 Days
Uranus ................13.2 Days .......................38 Days
Neptune ..............16.0 Days .......................44.5 Days
Pluto ....................18.5 Days .......................45 Days
Kuiper (65AU).....20 Days .........................58.7 Days (72AU)
Oort (52 KAU) ....23 Months ......................2.26 Years (50.8 KAU)
The values indicated for a and d are maximum values. Due to relativistic mass variation these values vary over appreciable ranges during interstellar travel, much less so for travel within the solar system.
2. Provides the design, development, and emplacement of a highly reliable asteroid/comet deflector/killer system. Detect each which is in a collision trajectory with an inhabited site and tug/deflect each into a catastrophe avoiding orbit to protect earth and each colonized planet/moon from hazardous impacts. For rapidly rotating objects the beam will be aimed at the pole creating a crude natural thrust engine from which the applied force will result in the most advantageous motion to us of the object. The beam (more than one if required) can be applied at the speed of light thus allowing us to divert "late" discoveries. A relay station may be required at Jupiter or Saturn to protect sites beyond these planets--40 to 80 minutes delay could be fatal.
3. The development of a reliable five billion year energy supply and delivery system to replace the fossil fuels which are limited in quantity and on which earth industry is now dependent. Fossil fuel reserves, if any by the time the system is operable, can be applied to the development of composite materials and the production of plastics.
4. Survival of earthling—especially human— species by distributing earthlings and earthling compatible life forms [minimum ecosystem sufficient for survival and colonization] throughout the solar system and the systems of nearby stars prior to an alien colonization of these areas. Also, it will become less likely that all humans will be exterminated by large catastrophic events, E.g., supernova: Explosion could annihilate all humans and many members of supporting species.
5. Accommodate the population explosion of humans. This may become a necessity more rapidly than we have assumed. Human population density and population density distribution drive the increase in difficulty of solving the technological aspects of the three more pressing and tightly coupled problems that we currently face:
1. Power generation and distribution.
2. Fresh water management and distribution
3. Pollution management and pollutant disposal.
6. Facilitation of pollution prevention and solar disposal of toxic wastes of all types(nuclear, biological weapons materials, chemical weapons and industrial waste products to name the more threatening).
7, Development and maintenance of solar orbiting laboratories, including a superconducting super-collider, well isolated from earth and in a naturally cryogenic environment. These laboratories require the use of extremely hazardous processes in necessarily extremely hazardous environments to “somewhat safely” develop nuclear fusion and genetic engineering techniques and genetically engineered products,which require complete isolation from normal human inhabited environments during development to minimize adverse consequences.
8. Development of technologies to obtain from nearby stellar systems material which may be different [or occur in different abundancy ratios] from that common to the solar planets. E.g., remnants from collisions involving black (burned out) carbon star dwarfs may include diamonds up to several hundreds of kilograms of mass; other objects may include rich deposits of elements essential to the technology development required to produce highly efficient photovoltaic cell performance, superconductivity, nuclear fusion and matter/antimatter engines and processes. It is difficult to specify the upper bound for the positive synergies that may evolve.
9, Emplacement, operation, support, and maintenance of human operated instruments in solar orbit at several times the distance of Pluto (or whatever sufficiently far from the sun turns out to be). This enhances our ability to measure extragalactic phenomena and high energy radiation sources as well as to locate earthlike (or other life supporting types of) planets in other stellar systems.
10 Mining of the gas giant planets for helium and hydrogen and robotic deep probing of their atmospheres to discover whether useful compounds or exotic lifeforms have formed under the high pressures there. Jupiter and Saturn may have large quantities of helium-3 which may jumpstart our fusion industry.
11. The advancement of technology and manufacturing processes, especially the spaceborne ones, including strong emphasis on the acceleration of our spacefaring expertise towards holding our own against any unwelcome nonearthling colonization of any Solar planet or any moon thereof.
12. Enhancing the development of high speed low cost (eventually) transoceanic/ transcontinental plasma propulsion based transportation for Earth's surface and atmosphere.
13. Initiate the development and evolution of technologies and design processes to begin intergalactic travel, black hole unzipping, galaxy configuration control, planet relocation, and universe expansion control.
MAJOR TECHNICAL RISKS
1. Containment of 388 billion square meters of receiver panels including the support structure within 5.94 billion kilograms of the mass budget (15.2 grams per square meter)---a major challenge.
Reduction Strategy:
a. Increase kw/m^2 capacity by investigating photovoltaic materials performance at high energy densities and high temperatures.
b. Vary the size of the small angle between the power beam and the receiver surface to optimize the effect of the cavity-like configuration of the "folded" panels.
c. Lower total energy required by going slower (e.g., 0.01 to 0.1 g accelerations).
d. Explore induction schemes to directly transform beamed energy to electric power for vehicle propulsion and operating power.
e. Determine the optimum heat exchanger including the operating medium to use cooling system heat to drive turbines for vehicle operating and thrust supplemental power.
2. Achieving beam collimation to maintain 90% of energy within a cylinder of 8-kilometer diameter over a distance of 5 light years.
Reduction Strategy:
a.Test and Analysis of proposed collimation schemes.
b.Verify quality of collimation at 40 AU.
3. Propellant collection at IV velocities greater than 0.00001c.
Reduction Strategy:
a. Test and Analysis of proposed methods.
b.Beam ionized particles collected from the stellar wind from star of origin.
4. Obstacle detection and avoidance at high IV velocities.
Reduction Strategy:
a. Test and Analysis of models of proposed approaches.
b. Onboard radar for IV velocities less than 0.02 c; time coded signals from the PBG for larger IV velocities.
5. Maximizing receiver panel energy density accommodation performance to partially abate risk number 1--may require several megawatts/square meter.
Reduction Strategy:
Test and Analysis of receiver panel configuration models in laboratories and explore performance of various new material combinations.
6. Protection of the PBG from solar mass ejection densities present at 0.05 or less AU.
Reduction Strategy:
Design shielding to protect against twice the known maximum mass ejection densities.
7. Production and management of the initial set of launch systems required to place the first PBGs in appropriate orbits.
Reduction Strategy:
Integrate energy collected from wind and photovoltaic units to produce hydrogen to use as fuel for the most efficient chemical combustion rocket engines.
MAJOR POLITICAL RISKS
1. Initiating and maintaining credible assurance that the destructive power of the collimated beams will not be used as a weapon.
2. Initiating and maintaining credible assurance that appropriate measures will be taken to protect earth's environment.
3. Initiating and maintaining credible assurance that the economic benefits will accrue equitably to each nation.
4. Initiating and maintaining credible assurance that the Earth Launch Facility will not pose hazardous conditions for any earth inhabitants.
5. Initiating and maintaining credible assurance that the technical and financial management of the project will be fair and reasonable.
6. Failing to start the program in a timely manner.
IMPLEMENTATION SPECULATION
By the time a preliminary design is mature, sufficient data will be available from the ongoing planetary astronomical observations and analysis to drive the decision for selecting the target stellar system to visit first. The assumption is that 10 light years is the limit in range due to our beam collimation skills. For now the Alpha Centauri system is the most likely target. The "stepping stone (star)" aspect of this concept is orders of magnitude more easily implemented where suitable platforms, planets, asteroids, and/or moons, are orbiting the stars in the systems to be visited. Such platforms are very convenient for colonization and for launching subsequent missions to more attractive locations further from the sun. Otherwise PBGs will have to be provided from a previously colonized system. Since each star is a gravity well, it is very likely that enough asteroid/comet material exists in orbit around the star to build suitable platforms where none have formed naturally. Early on we need to identify the minimum set of equipment and tools to transport to colonization targets. Design and process verification requires that robotic craft be sent to a target system to qualification test each facet of the concept not otherwise testable prior to putting humans at risk. This provides an opportunity to gather data about the target system useful for system design optimization and for colonization planning. Also, it will help us avoid infringing on the inalienable rights of any sentient critters already there as well as allow the establishment of communication with such cultures. The suggested implementation plan consists of four phases as follows:
PHASE I: PROGRAM FEASIBILITY VERIFICATION An Interstellar Transportation Consortium will be established and a Directorate will be appointed (elected) to direct the project including the approval of the completion of each major program milestone and the selection of the prime and integrating contractors. The goal is to have as many nations participate as are willing. Each major established space agency will be a voting member of the Directorate. Also, there will be a voting member representing each group of participating nations not currently part of a space agency. Each group will represent 10 nations. Each group will be formed by choice of its individual members and the rules for instructing how the group's vote is to be cast will be the responsibility of each such group. Political oversight of the project will be through the United Nations by whatever rules and agreements they are able to formulate. Each voting entity will be represented on the Security Team, which is responsible for the direction and control of the PBG energy beams and all safety issues attendant thereto.
A. Technical proof of concept:
1. Let contracts to begin manufacturing prototypes of components and major assemblies as required to provide proof of concept.
2. Conduct proof of concept testing.
B. Mutual Political Assurance that the Beam will not be used as a weapon and agreement on safeguards.
C . Planetwide agreement on financing of the project and mutual sharing of the economic benefits.
PHASE II: PRELIMINARY DESIGN
APPROVAL
A. Interstellar Transportation Consortium Directorate will conduct a Preliminary Design Review (PDR) and select from among competing designs and processes the ones, which best satisfy project goals and schedules.
B. The prime and integrating contractors will present the preferred design solutions including the attendant manufacturing processes. For equivalent competing designs the Directorate may require that each be presented to the extent the Directorate specifies. The formal PDR will amount to an endorsement of designs and processes that have been under tight scrutiny and close coordination throughout the development process during the proof of concept research and development activities.
PHASE III FINAL DESIGN ACCEPTANCE
The Directorate approves the resolution of any problems arising out of the PDR, accepts qualification test results and approves the final design.
PHASE IV: PRODUCTION AND LAUNCH.
A: Place EOF PBG in geosynchronous orbit
1. Manufacture and test major EOF PBG assemblies.
2. Unmanned launches for equipment and construction material.
3. Shuttle (or better) launches for assembly and test personnel.
B. Construct EOF Launch Facility.
C. Manufacture and test major EOF vehicle assemblies.
D. Perform final assembly and system integration of the EOF vehicle.
E. Manufacture and test major IV PBG assemblies.
F. Tow major IV PBG assemblies into orbit position, integrate and test.
G. Manufacture and test major IV assemblies.
H. Tow major IV assemblies into geosynchronous orbit, integrate, and test.
I. Begin trip to Proxima Centauri.
Note that the experience gained and the skills developed from this implementation will very likely naturally evolve into the capability to steer the solar system to locations more to our choosing. By placing an appropriately arranged network of enormous reflectors within a few hundredths of AUs of the sun we can cause it to propel itself in directions we choose. Gravitational coupling will bring each of the planets along as well. Further evolution of the concept may lead to matter/antimatter propulsion and the attendant intergalactic exploration prior to the year 3000 and the implanting of humans (or whatever they have become) within galaxy M31 in a mere 3 or 4 million years.
LOGISTICS AND ECONOMICS
The design disciplines and manufacturing processes exist today to produce this system with sufficient reliability (50 or more years life cycle for each major independent system) and maintainability (automated fault detection and isolation, built-in redundancies at modular level and the attendant autoreconfiguration, maximum use of interchangeable parts across all levels of assembly, etc.,) to generate relatively low life cycle costs. The 30-year view seems unaffordable; the 100-year view reverses the perception since the reusability of the system elements eventually brings the cost within the bounds of affordability. The design, development, test, and production of this system will be a global economic driver exceeding that of the combined effects of the jet transport and information service processing industries of the 20th century. Can such a rising tide fail to lift each boat? This will generate a veritable economic tsunami!
SERMON
Interstellar exploration and the colonization of extrasolar planets will greatly enhance human survivability. To delay development and operation of such a capability is unconscionable. It is the fate of sentient critters, if not their duty, to steer evolution in whichever ways seem most beneficial to them--they won't always guess right but this is preferable to leaving the steering to non-sentients whether biological or abiological. The universe abhors (does not permit) stasis and (eliminates) those who would promulgate it!
The earth, the cocoon for the pupa of humanity, is precariously attached to the solar system that is but a twig on the galactic tree. Now is the time for the imago to emerge and explore the tree----and eventually, its neighbors.
Shall our dreaming limit our vision of progress to having our descendants do less than engineer the expansion of the universe...with appropriate deference to the will and grace of God of course.
Feasibility Studies
The feasibility studies have been grouped into three time intervals spanning several years each to suggest priorities for focusing intensity. Most items will require continual revisiting to promote optimum evolution of the efficiencies the concept requires.
2006 thru 2009
For selecting photovoltaic assemblies [or other technology] capable of accommodating up to one GW/meter^2 energy density. Alternatively, determine the maximum capacity, input and output, of the photovoltaic technology in watts per square meter.
For determining whether 10e+20 tons or so of the more rare elements (indium, gallium, cesium, selenium, etc.,)which are required for developing the 50% efficient PV cells, can be obtained from the "surfaces" of Earth and the other likely colonization targets. Alternatively, determine what sort of efficiencies can be developed using silicon as the major constituent.
For developing convolutionary geometries and/or rotating surface schemes to accommodate reception of 10^18 watts over surfaces of manageable sizes.
For generating 10^18 watt collimated power beams and containing 90% of power within a cylinder of 4 KM radius over 5 light years. This includes selecting the optimum frequency with respect to receiving efficiency and collimation management. Alternatively, determine the maximum distance over which the above specified collimation can be achieved with current technology [and its continually evolving replacements] and evaluate the placing of collimation relay stations in galactic orbit between origin and destination stellar systems.
For generating megawatt mildly diverging auxiliary power beams for use by the ship to regain the main power beam after obstacle avoidance maneuvering.
For using PBG transmitted time coded references received at ship to determine position and velocity of objects hazardous to ships with velocities exceeding 0.1 c.
For anticipating the gravitational lensing effects on the power beam of objects lying in (or wandering into) the power beam between the IV and the PBG including determining the least mass for which gravitational lensing effects must be compensated.
For establishing criteria for determining whether the beam riding IV has lost its ability to regain the power beam due to the extent of the maneuver to avoid any hazardous object and develop the sophisticated algorithm required for the placing of the beam on the IV [in normal operation the IV steers itself to the beam center] and to restore both beam and IV to the desired trajectory toward the intended target stellar system.
For the inclusion of data content modulation in the power beam versus some auxilliary beam including the determination of minimum robustness of the communications receiver.
For producing hull skin and structure material at usable strength to weight ratios while accommodating enormous power density while keeping vehicle mass at a minimum. This includes evaluation of the ratio of heating energy to the sum of that reflected and that converted to electric power and the appropriate optimization.
For producing particle accelerator thrust engines including ones capable of working against 1 atmosphere external pressure [for earth launch vehicle].
For determining suitability of various molecules and the attendant processes for producing ions without separating the molecules into their constituent atoms [for ampere load management] determine optimum atomic/molecular weight versus interstellar availability versus ampere/volts trades in linac thruster design.
For determining optimum design of electrostatic grids in particle accelerators including material selection to minimize grid erosion and to facilitate the repair and/or replacement of grids.
For determining energy losses due to radiation from accelerating ions and for developing schemes to minimize such losses.
For implementation of beam pointing algorithm to compensate for exceeding galactic escape velocity [overcome time lapse between transmission and receipt of signals due to distance between ship and beam generating station] while directing the beam riding ship to the desired target.
For locating PBG near earth versus near mercury etc., Trade location cost versus available energy from various solar constant values as a function of orbit semi-major axes.
For determining environmental effects of discharging EOF ion exhaust in earth atmosphere including the radiation hazards, if any, from electron-ion recombination effects especially during and shortly after takeoff and low altitude portions of the flight trajectories.
For determining whether enough material can be collected from the solar wind to be used by the PBG to maintain its orbit position and station orientation.
For determining how to protect the PBG from solar mass ejections.
2010 thru 2013
For determining whether the amount of propellant material collectible from galactic orbit can support a 0.1 KG/sec propellant exhaust rate [q]. [If the Oort cloud outer radius is ~1.2 light years, and that is considered to be the solar system boundary with points on spheres of larger radii being considered intergalactic space and a similar cloud is assumed for the Centauri system and of proportionate radius, the probability of finding sufficient propellant between Sol and PC approaches 1. ] The 0.5c velocity vehicle will travel through a cylinder of length 150000 kilometers and diameter of 6 kilometers each second in intergalactic space. Also, determine the maximum vehicle velocity at which propellant collection is at all possible [including controlled magnetic field braking/acceleration of material prior to as well as during its capture].
For determining amount of reserve propellant required to maintain thrust when passing through unusually mass sparse volumes of space.
For determining whether propellant can be beamed in ionic form (solitons come to mind) from stations located near trip initiating stars to arrive at the ship at a few hundreths of light speed relative to the ship’s speed.
For beam rider earth launches versus existing launch systems including generating hydrogen fuel using solar energy.
For providing adequate protection from the hazards caused by encountering low energy photons by a ship traveling at near (e.g., 0.9 c) light speed.
2014 and beyond
For maximizing energy-produced-to-mass ratio of photovoltaic based power converters.
For generating plasma at up to 5 kilograms per second ion exhaust rates [q].
For establishing the practical upper limit for ion exhaust rates.
For prevention of hazards to earth environment from exhaust of
ions traveling at .99 to .999999 light speed from earth launch vehicle.
For designing launch facility to provide the necessary degree of isolation and/or protection from the hazards caused by the ion exhaust.
For developing the technology for maximizing the power density
capability of shipboard receiving designs.
For developing light weight photovoltaic materials and design configurations with 50 percent or more energy conversion capability.
For developing embryo sustaining and development canisters [growing a fertilized egg of a large mammal into a viable fetus in the uterus of a rabbit sized mammal or its test tube simulation] for the embryos of up to Clydesdale-horse-sized mammals to populate colonized systems with familiar protein sources.
Alternatively develop reversible miniaturization of large food source mammals. [versus supplying these items by freighter or from genetic engineering once the likelihood of colonization is firmly established.]
For evaluating energy loss from power beam due to ship propulsion ion exhaust through part of which the power beam must pass.
For designing the plasma thrust system to accommodate the need to spend the
appropriate amount of time in deceleration and acceleration while compensating for the sail effect (positive for acceleration and negative for deceleration).
For determining the most efficient method of accommodating Doppler shifts of the frequency of the main power beam due to the constantly changing velocity of the IV relative to the PBG.
For determining whether sufficient hydrogen or water exists in intragalactic space to support an exhaust rate of 5 grams per second and whether the thrust system must be designed to use a range of elements and molecules thus imposing constraints on the mass spectrograph. What is the upper limit of the propellant exhaust rate that can be supported?
For determining how to transfer 5 bars of the atmosphere from Venus to each Mars and the moon.
For sketches of major components click here:
mbmcneill7@comcast.net
