1. Nuclear and Plasma Space Propulsion

1. Nuclear and Plasma Space Propulsion - Chapter 1 NUCLEAR...

Info iconThis preview shows page 1. Sign up to view the full content.

View Full Document Right Arrow Icon
This is the end of the preview. Sign up to access the rest of the document.

Unformatted text preview: Chapter 1 NUCLEAR AND PLASMA SPACE PROPULSION © M. Ragheb 2/22/2009 1.1 INTRODUCTION In their role as stewards of life on Earth and perhaps in the whole known universe, humans have a duty to preserve and spread life. With their acquired intelligence, science and technology, it is their sacred destiny to preserve life with the equivalent of Noah’s Arks on both the Moon and Mars. Life can be subject to extinction on Earth either from within through volcanic eruptions or viral epidemics or from asteroid or comets impacts from space, as we know may have happened in the past. It is urgent to keep backup copies of life, like we keep for files on computers, on the moon and Mars protected from the possible unexpected calamities that could extinguish life on Earth. Fig. 1: The power requirements versus the length of mission stay times for space missions. Large amounts of chemical energy must be used in space travel to propel the space vehicle, especially out of the main pull of the Earth's gravity. The first stage of the Saturn V rocket used in the Moon missions Apollo program generated as much energy as 1 million automobile engines. The rocket engine as well as the propellant fuel must also be compact and lightweight, before the space vehicle can carry them. Fig. 2: Apollo 11 boot print on the moon, July 20, 1969. NASA Photograph. The power requirements versus the length of mission stay times away from Earth favor solar and nuclear energy means. As shown in Fig. 1, for large power needs, nuclear propulsion becomes the only alternative, particularly in the deep reaches of space where solar energy is not even available. 1.2 HUMAN DESTINY AND SPACE TRAVEL The American astronomer Frank Drake, an originator of the Search for Extra Terrestrial Intelligence (SETI) Project, suggested in 1960 an equation considering the probabilities of existence of intelligent life in the universe and estimating the possible number N of planets with Earth-like life with a technological civilization in the known universe as: N = R*.Pp.ne.Pl.Pi.Pc.L (1) where: R* is the number of stars systems, Pp is the probability of occurrence of stars with planets, ne is the fraction of planets with habitable environments, Pl is the probability that life has originated on a given planet, Pi is the probability of life evolving into intelligence on a given planet, Pc is the probability that the evolved creatures have the technology to send signals, L is the Longevity factor. Other factors could be added to this equation. For instance a factor possibly designated as the moon effect, Pm, can be added for the probability that a planet would possess an Earthmoon balance relationship like the existing one. It is suggested that the Earth was initially at half the size of the present planet, when a celestial object, collided with the Earth forming a double planet. A smaller body separated and orbited the larger one initially, then moved away with a decreasing orbital velocity forming the present-day moon. Maybe the Earth and Mars collided in the distant past ejecting the moon in the collision, with the Earth as the larger size object keeping the Mars’ water forming its oceans. It is surmised that the Pacific Ocean, may be a remnant of that collision. Regardless of its mode of formation, the gradually increasing drag of the moon's gravity slows down the rotational wind of the Earth, which would otherwise blow at an excess of 200 miles per hour. Such a high wind, like what exists on Mars, would make the Earth's surface uninhabitable by surface dwelling creatures. Interestingly, a future Mars colony may have to be built underground or carved into the sides of cliffs for this very reason. The moon tides also contribute to the molten state of the Earth's core in addition for radioactivity. The ensuing Earth's magnetic field is offering protection for life on Earth against the solar wind. A factor PJ can be added for the so-called Jupiter Effect. Weren't it for the fact that Jupiter is the size that it is, and has a circular orbit, Earth would not exist as far from the sun as it presently is. At 50,000 miles closer, it would be too warm and outside the livability zone. A few thousand miles farther, water would be in the frozen state like on Mars. Jupiter with its large gravitation also attracts asteroids and meteoroids from impacting Earth and causing frequent mass extinctions. Most galaxies possess periodical Gamma Ray Bursts (GRBs) that sterilize the whole galaxy with intense electromagnetic gamma radiation precluding any higher life forms. This would add another probability for the absence of gamma rays bursts: Pγ, and implies that most galaxies do not contain organic life. Most yellow dwarf stars like our sun, are large enough to emit solar flare of large proportions so as to reach Jupiter's orbit. Another probability Py, would account for the fact that our sun is a yellow dwarf of just the right size so as not to destroy the Earth with its 11 or 22 years cycles flares. Periodic mass extinctions are a fact of the fossil record. It is suggested that some of them were caused by cosmic collisions by comets or asteroids with the Earth. A factor Pcc for the probability of surviving such collisions can be added to the equation. Thus one can suggest a modified form of the previous equation as: N' = R*.Pp.ne.Pl.Pi.Pc.Pm.PJ.Pγ.Py.Pcc.L (2) An optimistic view would identify the factors in Eqn. 2 as possibilities rather than probabilities. In possibility theory, the AND logical gate implies taking the minimum of the possibilities rather than the product of the probabilities as in probability theory, yielding: N' = R*.n e .L.Min[π p ,π l ,π i ,π c ,π m ,π J ,π γ ,π y ,π cc ] (2)’ where the possibilities πi replace the probabilities Pi. The only factor that is well understood in Eqns. 1 and 2 is R*. It is thought to be equal to 100-400 billion star systems in our galaxy alone. The Terrestial Planet Finder (TPF) satellite experiment by detecting the light reflected from a distant rocky planet, while nullifying the light of the parent star, may gain information on the probability of occurrence of stars with planets; the Pp factor. At present, several planetary systems have been found. Geoffrey Marcey and Paul Butler have discovered two thirds of these extra-solar planets, from the University of California at Berkely. They developed a technique to detect planets based on the gravitational pull of planets on their own sun like stars, causing a Doppler's effect wobble in the frequency of light coming from the star. Jill Tarter describes the present state of knowledge: “The Drake equation is a wonderful way to organize our ignorance.” Regardless, our most important endeavor is in making the longevity factor L as large as possible, since humanoids have been around on Earth for only 125,000 years. As suggested by Achenbach and Essick: “It is not clear yet that a brain like ours is necessarily a long-term advantage. We make mistakes. We build bombs. We ravage our world, poison its water, foul its air.” The Optical Gravitational Lensing Experiment in 2006 detected the first evidence of a solar system about 5,000 light years away that contains two scaled down gas giant planets that are the same distance apart as Jupiter and Saturn are from our sun, leaving room for a possible planet like Earth. The detection occurred when the star orbited by the planets crossed in front of a star farther from Earth producing gravitational micro lensing. In such a situation the nearer star’s gravity magnifies the light shining from the farther star. The planets’ orbits of their parent star altered this magnification in a distinctive pattern. The two planets have masses that are about 71 percent of Jupiter’s mass and 90 percent of that of Saturn. The parent star is about 50 percent the mass of the sun. The astronomer Carl Sagan estimated that there are a million technological civilizations in our galaxy alone. Frank Drake offers the number 10,000. John Oro guesses the existence of 100 civilizations in the Milky Way galaxy. There is a possibility that their assessments are incorrect. What if N is equal to just 1 in the Drake's equation? This would mean, as suggested by Ben Zuckerman from UCLA, that we may well be alone in this galaxy, if not in this whole universe. If, as suggested here, extra probabilities or possibilities are added to the original equation, one should be amazed at how our existing form of life even exists, and that N' = 1 in Eqn. 2 is indeed a unique event. 1.3 FUTURE FATE OF EARTH Robert Frost, the poet wrote: “Some say the world will end in fire, others say in ice.” The unique event of life on Earth will not last eternally. Biological life on Earth depends on the sun which will not last forever. The solar constant is gradually increasing, with the sun becoming brighter and hotter and larger. As the temperature of the Earth reaches 140 degrees F, the Earth will start losing its water supply. The atmosphere will be 10 to 20 percent water vapor rising to the stratosphere. There, water would break down chemically into oxygen and hydrogen. The hydrogen will escape into outer space. The oceans could disappear in about 1.2 billion years, turning the Earth into a lunar landscape. Even sooner, warmer temperatures will cause the oceans to absorb a higher concentration of carbon dioxide, which is essential for plant life. In about 500 million years, plant life would disappear as well as all life forms depending on plants. If the life of the Earth has been 4.6 billion years so far, with a 1/2 billion years left, the Earth is indeed in its old age. As the sun exhausts its nuclear fuel and expands outwards in about 3.5 billion years, it will engulf with hot gases Mercury, Venus and the Earth. Instead of fire, Earth could suffer from ice. If the gravity of a passing star disrupts the orbit of Jupiter, this could disrupt the Earth's orbit sending it into the cold of deep space. We can modestly suggest, for life's future's perspective, that human knowledge about nuclear, plasma and radiation phenomena is necessary for our destiny as humans to help the living universe getting borne. Space exploration and eventually travel and colonization within and beyond the reaches of our solar system will depend on radioisotopes, nuclear and plasma energy means for propulsion and survival. Note that there is no solar radiation to depend on at the far reaches of the solar system and beyond it. Humans are bound to biologically engineer new forms of life adapted to the vacuum of space or on the surface of frozen moons, comets and asteroids. Such mobile life will free itself from the planets’ gravitational traps inhibiting its free movement. As Freeman Dyson suggests: “Perhaps our destiny is to be the midwives to help the living universe to be born. Once life escapes from this little planet, there'll be no stopping it.” 1.4 NUCLEAR ROCKET PROPULSION CONCEPTS The nuclear rocket involves a combination of the principles of rocketry and nuclear reactor technology. Most of them involve the delivery of energy as heat or kinetic energy to the rocket itself or to a working medium such as liquid hydrogen. The working medium is then expanded through a nozzle and accelerated to high ejection velocities reaching 6,000 to 10,000 m/sec. The heating of the gas is not achieved by chemical reactions like in chemical rockets, but from nuclear reactions including fission, radioactive decay, fusion, and miniature nuclear explosions. In chemical rockets, energy is obtained from the propellants themselves, whereas in nuclear rockets the energy source and the propellant are separate. Several concepts have been proposed: 1. Nuclear Thermal Propulsion (NPT): In this case a fission reactor produces the energy generated from the fission of uranium. This energy is transferred to liquid hydrogen as a working fluid. The reactor core operates at high temperature above 2,200 degrees Celsius. A diagram of a solid core reactor thermal system is shown in Fig. 3. Many concepts for both the power generation and the propulsion aspects are under consideration. These include solid liquid and gaseous fuel reactors and liquid metal and gas cooled reactors. Solid core reactors include pellet beds, particle beds, wire core, and foil reactors. Liquid cores include a droplet core and a liquid annulus core. Gaseous cores include an open cycle, a vapor core and “light-bulb” concepts. Thermal to electric conversion cycles include dynamic cycles: Potassium Rankine and Brayton, as well as static cycles: thermionic and HYTEC. 2. Nuclear Electric Propulsion (NEP): The nuclear electric power generated from the fission reaction or from the decay of radioisotopes is used to accelerate ions or other subatomic particles, which are ejected from the back of the rocket providing in a continuous low thrust. Such a system is shown in Fig. 4, together with the other components of the nuclear electric vehicle including the payload, shield, radiators, thrusters, power conversion, and power conditioning equipment. The propulsion concepts include steady state and pulsed electromagnetic engines, pulsed electrochemical and steady state electrostatic engines. Fig. 3: Solid Reactor Core for Nuclear Thermal Propulsion, NPT. Fig. 4: Schematic of Nuclear Electric Propulsion, NEP Vehicle and System. Fig. 5: Schematic of a nuclear fusion rocket. Fig. 6: Mars exploration modules based on External Pulsed Plasma Propulsion, EPPP for a Mars mission. 3. Nuclear Fusion Propulsion: In this case, nuclear fusion using charged particles fusion reactions such as the reaction: 2 1D + 2He3 −> 1H1 + 2He4, (3) would produce only charged particles whose kinetic energy can be directed by a magnetic field from a nozzle at the back of the engine. Figure 5 shows a schematic of a fusion propulsion system including a thermonuclear plasma enclosed in a magnetic mirror generated by the conducting magnet coils surrounding the plasma. 4. External Pulsed Plasma Propulsion, EPPP: This concept using miniature nuclear explosive charges has been explored in the past and designated as the Orion project. The charges are ejected in the back of the rocket, and their energy is transferred to spring loaded plates at the back of the rocket. Figure 6 shows schematics of such a concept studied for a Mars mission. 1.5 NUCLEAR ROCKET PROPELLANTS In chemical rockets, the same materials perform the functions of working medium and energy source. The energy content of the reactant is controlled by the strength of the chemical bond. It becomes a major consideration limiting the rockets specific impulse to 500 lbf.s/lbm for optimal combinations of hydrogen with oxidants such as ozone, fluorine and oxygen. The choice of a chemical propellant is restricted by the propulsion parameters. Propellant mixtures of low specific gravity are favored from this perspective. In nuclear fission rockets, the propellant coming in proximity with the fission fuel must exhibit a low absorption cross section for neutrons. Table 1 shows the absorption cross section of some possible propellants. Low neutron absorption eliminates lithium and boron. Helium and beryllium face cost and handling problems. Thus hydrogen appears as a superior choice for fission rocket propulsion. On the opposite end in fusion rockets, a significant amount of the energy may be: carried away by neutrons such in the DT fusion reaction: 2 1D + 1T3 −> 0n1 + 2He4, (4) In this case the neutrons carry 80 percent (14.06 MeV) of the energy release (17.6 MeV). Thus elements with a high absorption cross section such as lithium would be favored in this case. It would have to be introduced away from the plasma itself into the fusion products at the downstream end of the reaction zone. If it contaminates the plasma it would quench the fusion reaction through emission of bremsstrahlung x-ray radiation, and in the presence of a magnetic field in the form of synchrotron radiation. Table 1: Properties of some possible nuclear propellants. Propellant H He Li Be B C N O Thermal neutron absorption cross section (barns) 0.33 0.0008 71.0 0.005 750.0 0.0045 1.7 0.0006 Atomic Mass (amu) 1.0079 4.00260 6.941 9.01218 10.81 12.011 14.0067 15.9994 Hydrides can also be used. Water is one of them, but it dissociates into hydrogen and oxygen at high temperature exceeding 2,500 degrees Kelvin. In addition it is highly corrosive as high temperature steam. Other hydrocarbons can be used giving a dissociated molecular weight around 8 at high temperature and pressure. The nitrogen hydrides ammonia and hydrazine give dissociated molecular weights of about 10, but present a health hazard. For a trip to Mars, water stored under its surface as permafrost could be mined for the return trip in a nuclear rocket, and its use needs careful investigation. 1.6 ROCKET PARAMETERS Rocket propulsion combines the principles of mechanics, thermodynamics and in the present case, nuclear science. Propulsion is achieved by applying a force to a vehicle to accelerate it. Alternatively it involves the application of a steady velocity against a resisting force. The propulsive force is achieved by ejecting a propellant at high velocity creating thrust. The total impulse It is considered as the time integral of the thrust force F(t): t I t = ∫ F (t )dt (5) 0 The time t is the burning time of the rocket, and the thrust force F(t) is a function of time. In rocket engines, the propellant or working fluid is carried aboard the vehicle being propelled. Accordingly, the duration of the mission is limited by the mass of the propellant carried. This imposes a premium on the rocket's specific impulse Is defined as the ratio of the total impulse per unit weight w of the propellant: t Is = ∫ F (t )dt 0 w (6) where the total weight of the propellant in terms of the mass flow rate is given by: t w = g 0 ∫ m(t )dt 0 g0 = 980.66 m/sec2 or 32.16 ft/sec2, is the gravity acceleration at sea level. For constant thrust force F and propellant flow, this equation can be simplified as: Is = F .t F F I = = = t g0 m p g0 m w w (7) This equation identifies the specific impulse as the total impulse F.t, per unit weight of the propellant g0 mp. The specific impulse Is is also called the specific thrust since in fact it is the total thrust It per unit weight w of propellant. The units of the specific impulse in the Système International (SI) system of units is: m Newtons sec2 =sec . = m kg m kg . . 2 sec sec sec 2 sec kg. The effective exhaust velocity is the average equivalent velocity in m/sec, at which the propellant is ejected from the rocket. It is given by: veff = Is F = g0 m (8) The specific propellant consumption is the reciprocal of the specific impulse. It is the required propellant weight flow to produce a unit of thrust force in an equivalent rocket. Its units are kgs per kg.second. It is expressed in terms of the ratio of propellant flow rate to the thrust: Specific propellant consumption = 1 w mg 0 = = Is F F (9) The impulse to weight ratio of a complete propulsion system is defined as total impulse It divided by the initial vehicle weight or propellant loaded vehicle weight. A high value suggests an efficient design of the rocket. It is given by: It I st Is (10) = = m f g0 w0 (m f + m p ) g 0 + mg 0 ) ( t Where mf is the final mass of the rocket after exhausting its propellant, and mp is the propellant mass. The thrust to weight ratio describes the acceleration in multiples of the gravity acceleration that the engine is capable of giving to its own loaded propulsion system mass. Impulse to weight ratio = Thrust to weight ratio = F w (11) The propellant mass fraction is defined as: ς= mp m0 = m0 − m f m0 = mp m f + mp (12) This fraction describes the quality of the design. A value of 0.95 means that only 5 percent of the mass of the rocket is hardware that is used to contain and burn a larger mass of propellant. The final mass does not include non propulsion system components such as telemetry, communications and guidance instruments. The mass ratio of a rocket or a stage is defined as the ratio of the final mass after the propellant has been consumed to the initial mass of the rocket: Mass ratio = mf (13) m0 As an example, we consider a rocket with the parameters given in Table 2: Table 2: Typical Rocket parameters. Characteristic Initial mass m0 Final mass, mf Payload and structure Duration of operation, t Specific impulse of propellant, Is Value 2,000 1,300 1,100 30.0 2,400 Units kg kg kg sec 3 N-sec /kg.m, sec The mass ratio of the overall vehicle from Eqn. 13 is: Mass ratio of vehicle = 1300 / 2000 = 0.65 The mass ratio of the rocket system is: Mass ratio of rocket system = (1300-1100) / (2000-1100) = 200 / 900 = .222 The rocket propellant mass fraction is from Eqn. 12: ζ = (900-200) / 900 = 0.778 The propellant mass is: mp = 2000 - 1300 = 700 kg. The propellant mass flow rate is: (mp/t) = 700 / 30 = 23.3 kg/sec The thrust is : F = Is.(mp/t) = 2400x 23.3x9.80 = 540,800 Newtons. The thrust to weight ratio of the vehicle is: Initial F/w0 = 540,800 / (2000x9.80) = 28, Final value = 540,800 / (1300x9.80) = 43. The maximum acceleration of the vehicle is: amax = 43 / 9.80 = 421 m/sec2. The effective exhaust velocity becomes: veff = Isg0= 2400x9.80 = 23520 m/sec. Total impulse is: It = Isw = 2400x700x9.80 = 1,640,600 N.sec. The impulse to weight ratio is: It/w0 = 540,800 / [(2000-1100)x9.80] = 187. Rocket engines produce thrust by transforming a working fluid to a gas by subjecting it to high temperatures and then expelling it at high velocity through a nozzle. In chemical rocket systems, the propellants as fuel and oxidizer themselves provide the energy source, and are raised in temperature by the heat of combustion. In a nuclear rocket, the heat is supplied by a nuclear reactor, which heats the propellant that is being exhausted from the nozzle. Given an equivalent energy release to the propellants used in both the chemical and nuclear system, hydrogen if used in a nuclear rocket would provide 3 times the specific impulse generated in the chemical system. Table 3 shows the specific impulse advantage of different nuclear rocket concepts compared with chemical propulsion. Table 3: Comparison of Characteristics of Rocket Propulsion Systems. Concept Chemical-solid or liquid bipropellant Liquid monopropellant Solar heating Nuclear Solid Core Nerva Enabler Cermet Wire core Advanced Dumbo Specific Impulse [sec] 200-400 Mars trip duration [days] Working Fluid Fuel Temperature [oK] H2 and O2 N2H4 2773-4573 180-240 N2H4 N2H4 1273-1573 400-700 H2 - 1,573 Duplex UC-ZrC-C UO2-W UN-W UC-ZrC 2,270 2,700-3,300 825-850 925-1080 832 930 - 434 3,030 2,700-3,300 Pellet bed Particle bed Low pressure Foil reactor Nuclear Liquid Core Liquid annulus Droplet core Gaseous Core Open cycle Vapor core Lite bulb Electrothermal arc heating Electrostatic ion Magnetoplasma External Pulse Plasma Propulsion (EPPP) Fission Fission/Fusion, Fusion 998 1,000-1,200 1,050-1,210 990 1,600-2,000 1,500-3,000 5,200 1,280 1,870 400-2,000 UC-TaC UC-ZrC UC-ZrC UO2 434 3,100 3,000-3,500 3,000-3,600 2,700-3,400 3,000-5,000 5,000-7,000 200 60-80 310 U plasma UF4-HfC H2 6,000-8,000 7,200 5,773 4,00025,000 3,00015,000 Cs - H2 - 5,00010,000 100,000 - Fission plasma Fission/fusion, fusion plasmas A hydrogen-oxygen mixture propellant is normally selected for the upper stage chemical engines in planned space missions. Since in a nuclear rocket, energy is generated by the fission process, liquid hydrogen alone can be used. 1.7 SPACE REACTOR EXPERIMENTS: Named after the Kiwi, a flightless New Zealand bird, a reactor was built and operated as a rocket engine. It is shown while transported on its rail from the assembly building to the test cell in Fig. 7. The reactor was operated at high power at a predetermined temperature level and duration representative of an operational cycle. At the University of Florida's Innovative Nuclear Space Power and Propulsion Institute, research was being conducted on advanced reactor fuels for space propulsion. The research focused on interlocked wafers of tricarbide nuclear fuel consisting of Uranium, Zirconium, and Niobium Carbide. High quality solid solution tricarbides with less than 5 percent porosity have been produced. Optimum processing parameters for producing hypostoichiometric tricarbides are being identified. The high melting point, high power density, marked corrosion resistance of this fuel could yield significant improvements in thrust to weight and specific impulse over NERVA/Rover nuclear thermal rocket designs. Fig. 7: The Kiwi-A space reactor being tested at high power. Fig. 8: The KIWI B-4A Fuel Element Cluster. The Aerojet Company conducted the first nonnuclear demonstration of a Liquid Oxygen (LOX) augmented nuclear thermal rocket (LANTR). This idea could more than triple the thrustto-weight ratio of a nuclear thermal rocket via the injection and supersonic combustion of oxygen in the rocket's nozzle. Thrust augmentation of up to 44 percent is attained. Tests with inert nitrogen injection confirm that half the thrust increase is due to combustion of the oxygen. At NASA-Marshall research center and Los Alamos National Laboratory (LANL), research continues on the development of nuclear systems for electric and bimodal propulsion applications. Tests involving simulated fuel and heat pipe modules and reactor cores using high performance electric heaters, are being conducted. An entire 30-kW core has been successfully operated. Plans call for end-to-end demonstration of a simulated nuclear electric propulsion system at the Jet Propulsion Laboratory (JPL) using this core, a compact power converter, and a small ion thruster. Testing began of flight demonstration modules for a 300-kW reactor core. Sandia National Laboratory is contributing a flight experiment design study centering on this reactor configuration. Fig. 9: Nuclear Jet airplane engine at Idaho National Engineering Laboratory (INEL). Fig. 10: Comparison of Chemical and Nuclear systems for a Mars mission. 1.8 MARS MISSION PROPULSION REQUIREMENTS The true potential of a nuclear rocket is not just for providing power for observation satellites and anti ballistic weapon systems, but for a possible space mission to Mars. The higher specific impulse of the nuclear rocket can reduce the mission time for a Mars mission from about a year for a chemical rocket, to about 2 weeks in the case of a fusion rocket. This may be crucial to avoid the effects of space radiation from solar flares on the astronauts, as well as avoiding the effects of gravity's absence on the muscular bone, and other bodily functions from exposure to space radiation and solar flares in long duration space missions. Figure 10 compares the chemical and nuclear fission vehicles required to perform a manned Mars exploration mission. Assuming that the space vehicle has been assembled in an Earth orbit, with the components supplied by a space transport vehicle, or reusable rockets, the all-chemical vehicle would have an initial weight in Earth orbit of almost 10 million pounds. The nuclear vehicle weight would be about 1/10 this value, at about 950 thousand pounds. The weight advantage is here clear. A nuclear rocket would be crucial for the return of the astronauts. The USA NERVA reactor as well as Russian designs used U235 as the fuel. New fuels consisting of tricarbide fuel: (U235, Zr, Nb)C. The use of Pu239 is precluded by United Nations agreements on the use of space. The use of a nuclear rocket cannot be used for landing and return from Mars. Because of its radioactive exhaust, and the added need for surrounding, rather than just shadow shielding of the crew, the landing and return must use chemical rockets, with the nuclear rocket left in orbit around Mars. This is necessary, since the effective dose rate from an unshielded NERVA engine after being fired can be in the range of 10,000 rem/hr, so that the crew cannot stay close to it, should it be landed on Mars. As an illustration, the fission product activity produced from a run lasting 1,000 seconds from a 2,000 MWth nuclear rocket would produce more than 109 Curies (Ci) of fission products, which is 1/10 what is produced over two year operational period for a typical land based 3,411 MWth nuclear power plant. The Orbitech company has been developing in-situ resource utilization systems to exploit the Martian atmosphere for ground transportation, flight propulsion, and power. Solid CO and C are used as fuels in hybrid rocket propulsion systems. Small-scale solid CO/O2 hybrid motors, cryogenic solid hybrid rocket engines, vortex combustion ramjets, scramjets, and solid oxygen/liquid hydrogen hybrid engines are being pursued. Because of planetary alignments a window of opportunity for a trip to Mars opens every 26 months, with some windows being better than others. The year 2016 would offer a good window. NASA's Johnson Space Center estimates the cost of a mission including 3 trips to Mars at $50 billion. A scaled down approach could be done for 20-30 billion in 2000 dollars. On Mars, nuclear power would be needed. Because of dust storms and high wind speeds, a Mars colony would have to be sheltered underground, and need a reliable power supply for heat, transportation, food production, water supply, communications and other life supporting measures. The environment on Mars is very harsh. Temperatures average at below 273 degrees K, and are at 148 degrees K at the Polar Regions. The climate is dry and hostile, threatening the astronauts at every turn. Providing energy, particularly heating for the astronauts cannot depend on solar energy or on radioisotope generators, and needs a nuclear reactor source. A mission composed of 4 astronauts would need a power supply of about 140 kWe. Most radioisotope generators have used plutonium238, and assuming a dynamic conversion system's efficiency of 30 percent, the thermal energy needed for the astronauts is 140 x (100 / 30) = 466.66 kWth. One needs about 1.8 kg of Pu238 per kWth produced. Thus one needs: 1.8 x 466.66 = 840 kgs of Pu238. This amount is beyond any possible existing supply, and suggests that such a mission, for reliability reasons, would require at least two nuclear reactors producing a thermal power of 0.5 MWth each, for a total of 1 MWth of power. During the Martian day, three solar power systems at 10 kWe each, may supplements their needs. It will take at least a decade of research and development, with an expense of at least $50 billion to prepare for a Mars mission. NASA has been lately trying a strategy of "faster, cheaper, better," in its exploration of Mars, leading to about a 2 out of 3 as a success rate. With manned space mission, a higher degree of reliability will be needed. 1.9 EXTERNAL PULSED PLASMA PROPULSION INTRODUCTION This is a nuclear propulsion concept generating its thrust with plasma waves generated from a series of miniature supercritical fission or fusion pulses. The intense plasma wave energy transfers its momentum into vehicle acceleration that can be withstood by the structure of the vehicle and its crew. Very high specific impulses and thrust to weight ratios can be obtained by this approach, which other technologies cannot obtain. Their appeal also stems from their low costs and reusability. They offer fast interplanetary transit times, safety and reliability, and do not require major technological breakthroughs. This could be the only realistic approach available with present day technology for a Mars mission in the twenty first century. THE ORION PROJECT The USA Air Force pursued this project on a classified basis between 1958 and 1965. The proposed space vehicles would be 10-30 meters in diameter, as shown in Fig. 11, since the performance tended to increase proportionally to the diameter of the lower pusher plate. This is due to the higher specific yields or burnup fractions which increase with the size of the pulse units, as well as the wider solid angle intercepting the plasma from a larger plate, at the minimum standoff distance between the plate and the point of detonation. This distance is determined by material strength and materials ablation considerations. The ablative material pusher plate would absorb the impact and thermal shocks. The effort was not continued for political reasons. However it has been established that a space vehicle with high thrusts of 1-10 g accelerations, high specific impulse in the range of 10,000 secs can be built. The yield of the pulse units were in the range of 0.01 kT, and the repetition rate was in the range of 0.1-1 pulses per second. The standoff distance ranged from 100-1,000 feet. GOVERNING RELATIONS Consider the masses of the pulse unit device and the payload as: Pulse unit device mass Payload mass = md, = mp, As well as their respective velocities as: Pulse unit device velocity Payload velocity = vd, = vp. Fig. 11: An External Pulsed Plasma Propulsion, EPPP Space Vehicle. Assuming that all the energy of the pulse device, E is released in the form of kinetic energy, applying conservation of energy, then: E = ½ mdvd2 (14) The fraction of solid angle intercepted by the pusher plate in spherical coordinates is: θ f = R 2π θ ∫ dV ∫ ∫ ∫ r = 0 V = sin θdrdθdφ 0 0 0 4πR 3 3 R = 2 2π θ 0 0 0 3 2 ∫ r dr ∫ dφ ∫ sin θdθ 4πR 3 (15) 1 (1 − cosθ ) 2 The fraction of energy transferred to the pusher plate becomes: E’ = ½ f mdvd2 = 1/4 (1 – cos θ ) mdvd2 (14) Fig. 12: Geometry for the External Pulsed Plasma Propulsion, EPPP Rocket. Applying conservation of momentum yields: mp vp = md vd, from which: (15) vd = (mp / md ) vd Eliminating vd from Eqn. 14, yields: 2 2 1 ⎛ mp ⎞ 1 mp 2 fE = md ⎜ vp ⎟ = vp 2 ⎜ md ⎟ 2 md ⎝ ⎠ (16) One can thus deduce the payload velocity as: vp = 1 1 1 (2 Efmd ) 2 = 1 [E (1 − cosθ )md ]2 mp mp (17) and the device’s particle velocity as: 1 1 ⎡ ⎤2 1 (2 Efmd ) 2 = ⎢ E (1 − cosθ ) ⎥ vd = md md ⎣ ⎦ (18) Considering that the device’s plasma collides with the pusher with a velocity vd, and is reflected with a velocity in the opposite direction (-e.vd), where e is the collision elastic parameter, the change in momentum will be: d (mvd ) = m[v d − (−evd )] = mv d (1 + e) (19) The specific impulse in this situation can be written as: Is = ∫ Fdt mg 0 d = ∫ dt (mv mg 0 d )dt = ∫ d (mv mg 0 d ) (20) Substituting from Eqn. 19 into Eqn. 20 we get: Is = v d (1 + e) g0 (21) Substituting for vd from Eqn. 18, we get: (1 + e) ⎡ E ⎤ Is = ⎢(1 − cosθ ) ⎥ g0 ⎣ md ⎦ 1/ 2 (22) For an elastic collision, where the expanding plasma loses all its momentum to the pusher plate, e = 1, and: 2 Is = g0 ⎡ E ⎤ ⎢(1 − cosθ ) ⎥ md ⎦ ⎣ 1/ 2 (22)’ This equation shows that the specific impulse will be proportional to the specific yield of the device (E / md), and the subtended solid angle. The use of a fusion component would maximize this ratio. Devices where the energy is collimated through this solid angle, where the pusher plate would subtend most of the released energy, would be more effective than the spherically symmetric ones. DIRECTED ENERGY PULSE UNITS It is possible to direct the energy of a nuclear device through a chosen solid angle instead of distributing all its energy into a 4π solid angle using asymmetric burns. The energy from a nuclear device is channeled through a radiation case containing a channel filler to generate a plasma that transfers the energy to a propellant plate. Initially a low Z material was used in the Orion project. It was substituted for by a high Z element. This results in a dense plasma (high Z case) at a relatively low velocity at a wide angle, instead of directing a lower density plasma (low Z case) at a higher velocity and a narrower angle as used in the Strategic Defense Initiative (SDI) or Star Wars Project directed energy project designated as Casaba-Howitzer.. A schematic of a pulse unit for a 10 meters in diameter Orion vehicle is shown in Fig. 13. It would yield about 1 kT of energy, and weigh 311 lbs. About 2,000 to 3,000 charges would be needed for a return trip to Mars. The initial burst of energy is confined by the radiation case and channeled toward the propellant slab. Fig. 13: Design of a directed energy pulse unit with heavy element pusher plate. FREE EXPANSSION OF A PLASMA IN VACUUM The free expansion of a gas in a vacuum results in the propellant disc expanding in an asymmetric expansion fashion. Since the plasma fluid would have a larger pressure gradient in the axial direction of the plate, it will expand into the shape of a cylinder as shown in Fig. 14. Interestingly, an inverse process would also occur: The free expansion of a cylinder would result in a plate shape. Under asymmetric free expansion a pancake shaped plasma would expand into the shape of a melon, and a melon shaped plasma would expand into the shape of a pancake. The free expansion yields an expanded diameter to length ratio inversely proportional to the square root of the initial diameter to length ratio. ⎛D⎞ ⎜ ⎟α ⎝ L ⎠1 1 ⎛D⎞ ⎜ ⎟ ⎝ L ⎠0 For instance, starting with a plate with an initial diameter to length ratio: ⎛D⎞ ⎜ ⎟ = 4, ⎝ L ⎠0 (23) would result in a cylinder of diameter to length ratio of: 1 1 1 ⎛D⎞ = = ⎜ ⎟ = 4 2 ⎝ L ⎠1 ⎛D⎞ ⎜ ⎟ ⎝ L ⎠0 Thus starting from a flat plate can yield through asymmetric free expansion a plasma jet that is collimated within about 20 degrees. Fig. 14: Free asymmetrical expansion of pancake into cigar shaped plasma, and the inversion of a cylinder into a pancake shape. VARIABLE DENSITY PUSHER PLATE A further refinement in time shaping the pulse delivered to the Orion vehicle shock absorber can be achieved by controlling the distribution of density of the expanding plate. The load on the pusher is governed by the local density of the propellant plasma multiplied by the square of its velocity. A variable density in the plate can yield a softer ride and more effective horse power. In this case where an initial lower density is used in the back of the plate and a higher density in its front, or vice versa, the pressure pulse can be spread out over time or contracted, mitigating the shock to the space ship. This approach to pulse shaping is crucial for a viable propulsion system, instead of just having a rapid rise in pressure followed by an exponential decay. PRACTICAL CONSIDERATIONS As initially considered in the ORION project, the vehicle would be launched from the Earth’s surface. The release of radioactivity in the atmosphere was an unacceptable alternative at the time, and still remains so. However, if the components can be launched with a transport vehicle to low Earth orbit and assembled there, these objections disappear. The space environment is already extremely harsh in terms of radiation. It has more background radiation in the form of gamma rays than the small pulse units would produce. In a matter of 24 Earth hours, the resulting ionized mass would dissipate in the background space plasma density. The exhaust particles velocities would exceed the Earth’s escape velocity and even the solar escape velocity, resulting in no residue or permanent contamination above the level caused by the natural radiation from the sun. This technology is immediately available for space missions. There is no guarantee that other technologies such as fusion propulsion, matter/antimatter and beamed-energy sails that are under study will be available during the first half of the twenty-first century. Fusion must await the demonstration of a system possessing sufficient energy gains for commercial and space applications. Matter/antimatter has low propulsion efficiency and a prohibitive cost of the possible production and storage methods. Beamed energy would require tremendous investments in ground and space based infrastructure. The need for high power densities for space missions favors nuclear energy sources. Solid core nuclear thermal, gas core, and electrical nuclear propulsion systems have problems with the constraint of the need of containment of a heated gas, which restricts its specific impulse values. External pulse systems possess higher temperature limits and lower inert masses and circumvent that limitation. Several methods of external momentum coupling have been investigated other than the standard pusher plate. These include a combined magnetic field and pusher plate, a rotating cable pusher, and a large lightweight sail. Because the reaction is external to the material walls of the vehicle, the system’s operation is independent of the reaction rate, pressure temperature and the fuel characteristics. The physics of fission in a vacuum are simple where a shell of ionized gas with extremely large radial velocities is produced. It is also recognized that common materials can withstand an intense nuclear damage environment over short intervals of time in the nanoseconds range. The acceleration of the ship is only limited by human and equipment tolerances. Imparting high thrust for short periods of time results in fast and efficient trajectories. Current research emphasizes low ablation pusher plate designs, low energy pulse unit yields, and dedicated space operation out of the Earth’s atmosphere. The overall advantage is that this approach can yield space vehicle for a Mars mission of duration of just 1-3 months. This should be compared to the mission time of about 25 months with chemical or other propulsion technologies. The latter technologies favor Hohmann type transfers into very slow heliocentric orbital trajectories; which narrows the available trajectories for return and necessitates long stays on the Mars surface waiting for the occurrence of favorable return windows. This stay would be in an extremely hostile environment with 560 days surface stays and 170-200 days transit times. It would also provide more flexible return windows and eliminate the need for long stay times in the vicinity of Mars, where the astronauts’ bodies would be ravaged by the effects of a long period of weightlessness and high space radiation, in addition to the lurking deadly danger of unforecasted solar flares. Short duration missions on Mars provide by External Plasma Pulse Propulsion would also be associated with lower overall mission costs. Longer missions translate into a need for larger payloads and expandables that need to be launched into space at high cost. The specific impulse of nuclear thermal systems is in the range of 900 sec, which is about twice those of chemical propulsion systems in the range of 450 sec. The main advantage here is the reduction of the vehicle mass in low Earth orbit, thus reducing the number of heavy lift vehicle launches. External Pulse Plasma Propulsion is distinguished by specific impulses in the range of 5,000-10,000secs. Even higher specific impulses of 100,000 secs can be achieved with larger vehicles, and more energetic detonations using fission/fusion and fusion sources. These can open up the whole solar system for human exploration and colonization. 1.10 DAEDALUS SPACE SHIP PROJECT An interstellar space ship study was conducted over the period 1973-1978 by the British Interplanetary Society, designated as the Daedalus Project. A target was used as the Barnard’s Star system, a red dwarf at a distance of 5.91 light years from the sun that was sought at the time to be possibly orbited by planets. To reach its destination within 50 years, the ship had to cruise at 12 percent of the speed of light at 36,000 km/sec. A fusion propulsion system was selected using the D-He3 fusion reaction producing H and He4 charged particles that could be diverted using magnetic fields was selected rather than a fission system like the Orion project. The He3 isotope would have to be bred from lithium into tritium that would decay into He3 on Earth, or alternatively mined from the surface of the moon whose dust in rich in He3 deposited by the solar wind. Fig. 15: Conceptual design of the Daedalus space ship. Source: British Interplanetary Society. Fig. 16: NASA's Hubble Space Telescope was used to take the first visible-light snapshot of a planet orbiting another star. The images show the planet, named Fomalhaut b, as a tiny point source of light orbiting the nearby, bright southern star Fomalhaut, located 25 lightyears away in the constellation Piscis Australis. A large debris disk about 21.5 billion miles across surrounds the star. Fomalhaut b is orbiting 1.8 billion miles inside the disk's sharp inner edge. In an inertial confinement fusion system, fusion pellets would be irradiated by electron or laser beams. The charged particles products would be channeled by a magnetic field as a hot plasma out of a nozzle to provide the required thrust. A repetition rate of 250 pellets/second and the use of a two stage system would attain the cruising speed within a 4 years acceleration period. The space ship would be assembled in Earth orbit with a weight of 54,000 tons of fuel and a 500 tons scientific payload. The first stage would be fired over two years to attain 7.1 percent of the speed of light then jettisoned. The second stage would fire for the next 1.8 years for a 46 years cruise to the Barnard’s Star. About 18 probes powered by ion drives would be used to investigate the star and its planets. A 50 ton, 7 mm thick disc of high strength beryllium metal would be used to shield the payload bay from collisions with space dust and meteoroids on the flight. An artificially generated cloud of particles 200 km ahead of the vehicle would disperse larger particles as it reached the planetary system of the target star. It would take 12 years for the radio signals from the probe to reach Earth. Accordingly, the probe must be self autonomous using artificial intelligence controls. The first step for such a mission in the 21st century has been taken by NASA in its Deep Space 1 probe. 1.11 DEVELOPMENTS PROPULSION IN PLASMA, PHOTONIC AND LASER Magneto Plasma Dynamic (MPD) thrusters are being developed using a 250-kJ capacitor bank and pulse-forming networks at the NASA Glenn Laboratory. A high power steady vacuum facility is readied for for long duration MPD thruster tests at power levels up to 1.5 MW. Ion acceleration and heating methods for advanced plasma propulsion are being pursued at Princeton's Electric propulsion and Plasma Dynamics Laboratory. A coherent ion acceleration mechanism depending upon the nonlinear interaction of a magnetized ion with multiple electrostatic waves is being researched; at least two of which differ in frequency by an integer multiple of the cyclotron frequency. The ions need not be in resonance and be coherently accelerated with an arbitrary low initial energy. Magnetic field expansion for mini magnetospheric plasma propulsion is being pursued at the University of Washington and NASA-Marshall research center. This concept generates a magnetically confined plasma bubble that achieves thrust through interaction by interaction with flowing charged particles in the solar wind. Energy Sciences Laboratories (ESLI) constructed a microtruss fabric, from carbon fibers and whiskers. This thick porous material has applications in solar photonic sails. Areal densities of 1-10 gm/m2 were achieved, with demonstration of the elastic self deployment of these structures after stowage. Microwaves and lasers were used to impart momentum to small carbon sails at the Jet propulsion Laboratory (JPL). The sails were accelerated at several gs up to a height of 2 ft with a 10 kW microwave beam. Laser experiments at Wright Patterson Air Force Base (WPAFB) demonstrated horizontal deflection of a pendulum-mounted sail with laser power ranging from 7.9 kW to 13.9 kW. At the White Sands Missile Range, a thrust stand was used to perform static thrust measurements of Lightcraft models at different distances up to 120 ft from a 10 kW laser source. Laser to air energy coupling is being studied to increase the launches to several thousand feet. REFERENCES 1. Gary L. Bennett, “Nuclear Technology on Mars,” Book Review, Nuclear News, October 2000. 2. Gregory Benford, “The Martian Race,” Warner Books, 1999. 3. George Schmidt, “Nuclear and Future Flight Propulsion,” Aerospace America, December 2000. 4. Concetta G. Capoen, “United States Aerospace Nuclear Power Programs, A Compendium,” U. S. Atomic Energy Commission, 1961. 5. J. S. Levine, “Terraforming Earth and Mars,” in: Mars: Past, Present and Future, E. B. Pritchard, Ed., Progress in Astronautics and Aeronautics, Vol. 145, 1992. 6. D. A. Poston, “The Heatpipe-Operated Mars Exploration Reactor (HOMER), CP552, Space Technology and Applications International Frum-2001, M. S. El-Genk, Ed., American Institute of Physics, 2001. 7. J. S. Clark, “Nuclear Propulsion Technology Development: A Joint NASA/Department of Energy Project,” “Mars: Past, Present, and Future,” E B Pritchard, Ed., Progress in Astronautics and Aeronautics, A R. Seebass, Ed., Vol. 145, 225-237,1992. 8. G. P.Sutton, “Rocket Propulsion Elements, An Introduction to the Engineering of Rockets,” John Wiley and Sons, 1986. 9. E. M. Goodger, “Principles of Spaceflight Propulsion,” International Series of Monographs in Aeronautics and Astronautics, Division III. Propulsion Systems Including Fuels, Vol. 6, P{ergamon Press, 1970. 10. J. A. Bonometti, P. J. Morton and G. R. Schmidt, “External pulsed Plasma Propulsion and its Potential for the Near Future,” CP504, Space Technology and Applications International Forum-2000, M. S. El-Genk, Ed., American Institute of Physics, p. 1236, 2000. 11. George Dyson, “Project Orion, the True Story of the Atomic Spaceship,” Henry Holt and Company, New York, 2002. ...
View Full Document

This note was uploaded on 06/16/2010 for the course NPRE 402 taught by Professor Ragheb during the Spring '08 term at University of Illinois at Urbana–Champaign.

Ask a homework question - tutors are online