Emerginng technologies in space vechicle propulsion

PROPULSION SYSTEM
Introduction
In the design of most spacecraft, the propulsion subsystem is one of the key design considerations, especially when considering that the propulsion subsystem is many times the single most massive component of the spacecraft, comprising as much as 90% or more of the total spacecraft mass. Because of this, the propulsion subsystem performance affects both mission design and payload mass. There are a large number of different propulsion system types that are in use today or that have been proposed, and they cover a wide range of capabilities. The goal of this report is not to provide a detailed analysis of each type of propulsion type, but to classify and characterize various propulsion subsystem types and to assemble a database to provide the information necessary for the design of a new spacecraft's propulsion subsystem. The Propulsion System Database is available on disk for both the PC and the Macintosh.
Propulsion System Types
There are several rocket propulsion system types in common usage. All share the common characteristic of providing the propulsive force by the momentum of ejected matter, the propellant. The major difference in various propulsion systems lies in the method of supplying energy to the propellant. Energy is added to increase the kinetic energy of the propellant. This increase in energy results in an increase in velocity of the propellants being expelled, meaning an increase in propellant momentum. Since momentum is conserved for the system, the vehicle momentum in the opposite direction must also increase. Energy for the propellants can be supplied in a number of ways, including chemical combustion, nuclear reaction, solar radiation, and others. These different energy sources can be used to categorize various propulsion systems. Additionally, the function of the enginemay be used for characterization. These functions include apogee or perigee kick motors, attitude control, station keeping, precision pointing, reboost, landing, retrorockets for deorbit, and others. To help give the designer a clearer perception of the choices available when developing a propulsion system, the following page contains a schematic breakdown of the major propulsion system types organized in a tree form. Three major types of systems are included in the tree; Chemical , Non-Chemical and Exotic/Theoretical. Each of these categories is then broken down into more descriptive subcategories, showing the relationships between the various propulsion systems. The specific impulse and thrust range figures shown are only ranges of performance which categorize the specific propulsion system type. All the elements with an * next to them are described in more detail in the following text.
CHEMICAL PROPULSION SYSTEMS
Chemical rockets are the oldest form of rocket propulsion, invented over eight centuries ago by the Chinese for use in fireworks. Chemical rockets utilize chemical combustion to supply energy to the propellants. This combustion is typically the high pressure reaction between a fuel and an oxidizer. The combustion supplies heat energy to the combustion products, raising their temperature. To increase the combustion product momentum, a converging-diverging nozzle is used to expand the gases. Chemical propellant rockets can be grouped based on the physical state (solid, liquid, or gaseous) of their propellants. In the next four sections solid, liquid, gaseous, and hybrid propellant chemical systems are described.
Solid Fuel Rockets
Solid rocket motors are very common in American space operations. They are the simplest type of rocket engine as they have very few moving parts. The nozzle is connected directly to the propellant tank, with no pumps or other machinery between them. The propellant tank contains a mixture of fuel and oxidizer, which does not react until the engine has been ignited. Once ignited, the propellant burns continuously until depleted. The shape of the propellant is very important to the engine performance. A large burning area corresponds to a higher thrust than a smaller burning area. The propellant's shape may be tapered to provide periods of higher and lower thrust, according to the mission plan. The performance of the solid rocket is often measured by total impulse, which is the area under the Thrust vs Time curve. Solid rockets are often used when high reliability or long term storage of the motor is necessary. The specific impulses of typical solid rocket motors range from about 200 to 300 seconds, or only about 50%-75% as high as most liquid propellant engines. Table 8.3.1 contains theoretical performance numbers for a some common solid propellant combination. The primary advantages of solid engines are that they are the oldest rocket system, and thus are a well known and proven technology. They are easily sized, simple to operate, and reliable. On the other hand, the are not restartable, are dangerous and expensive to make, non or only partially throttleable, offer low specific impulse, and give off toxic emissions including chloro-fluorocarbons.
Liquid Fuel Rockets
Liquid fuel rockets, as with other chemical systems, utilize a chemical reaction to supply energy to the working fluid, the gaseous chemical reaction product, to obtain thrust. Liquid fuel rockets may be classified in several ways. One method to classify liquid rockets is by the number of chemicals involved in the reaction, either monopropellant, bipropellant, or multipropellant. Liquids may also be grouped as either storable or cryogenic.
Monopropellants are those chemicals which may be used alone to obtain a chemical reaction. These may be a mixture of several compounds or a single chemical. In general they are decomposed by heating or exposing to an appropriate catalyst, thus doing away with the typical ignitors necessary for other liquid propellants. This makes them ideal for systems needing high reliability and simplicity, such as attitude and trajectory control jets. They also have the advantage of only needing to store one propellant. Along with these advantages come the disadvantages that they usually have lower specific impulses than many bipropellants. They are also generally highly toxic substances involving special handling and storage. Some monopropellants which have been used include hydrogen peroxide, ethylene oxide, and nitromethane. By far the most common monopropellant is hydrazine, which offers a specific impulse of about 245.9 seconds at a chamber pressure of 1000 psi and expansion to vacuum.
Bipropellants are the most common type of liquid fuel rocket. As the name suggests, bipropellants use two components as reactants, an oxidizer and a fuel. In general, bipropellants the use of some type of ignitor to initiate the chemical combustion. However, there are certain combinations of propellants which will react upon contact. Such propellants are known as hypergolic propellants. Table 8.3.2 contains a listing of commonly used bipropellant combinations, along with some interesting theoretical combinations. Liquid bipropellants offer the highest specific impulses of all chemical systems at the cost of more complexity than monopropellant systems.
Existing liquid chemical propellant types can also be divided into two main classes, cryogenic and storable. Cryogenic propellants are gases which have been liquified under very low temperatures thus they must be refrigerated to remain in liquid form. Because of their low boiling points provisions must be made to account for boil-off losses during long space missions and to vent this boil-off, possibly for attitude or trajectory control as a gas jet which will be discussed later. Storable liquid propellant components have relatively higher boiling points and so do not require cryogenic refrigeration. These categories can be further divided into Earth storable and space storable. Earth storable propellants are liquids at Earth ambient temperatures and so are relatively easy to handle. Space storable propellants include those that are Earth storable and those which must be refrigerated on Earth but will remain liquid in space without refrigeration. Storable propellants are more suitable for missions in which the propellants must be held in reserve for long periods. Additional considerations in the design of liquid chemical rocket engines are the propellant feeding and ignition systems. Propellants may be fed into the combustion chamber by pressure alone or by some type pump. Pressure fed systems offer increased reliability and simplicity but not all propellant combinations can be pressure fed. Pump fed systems offer slightly higher engine performance at the cost of some reliability. They operate with a much lower propellant tank pressure so that they can make use of light-weight, thin-walled tanks and eliminate the requirement for a pressurization system. Calculations have shown that pump fed systems provide a greater payload than pressure fed systems for missions requiring more than 450 kg of propellant. Ignition systems are required for most propellant combinations and they add complexity and reliability concerns to the design of a spacecraft. Hypergolic propellants ignite spontaneously when they come in contact with each other and so do not require a separate ignition system. Monopropellants frequently utilize a solid catalyst to release the chemical energy of the propellant and so these systems also do not require separate ignition systems. Liquid propellants have the advantages of being a well known and tested technology, versatile and reliable. The main disadvantages are a maximum specific impulse of approximately 500 seconds, the need for complicated plumbing, and possible toxic emissions. An alternative to the oxidizer/fuel type of liquid engine makes use of the energy released from the combination of free radical species. The recombination of hydrogen radicals to form H2 has a theoretical specific impulse of 2,130 seconds, but no working versions of radical hydrogen engines have been tested.
Gas Jets
Rockets may use a compressed gas as the propellant. Typically these systems are primarily used for station-keeping and attitude control purposes. Gaseous rockets can be divided into two categories, cold (inert) gas jets and warm (heated) gas jets. Both systems expand high pressure compressed gases through a supersonic nozzle. The most common type of gas jet is the cold gas jet, and systems such as this have been used for periods up to three years on some spacecraft. Warm gas jets are similar to cold gas jets except that a heat source such as an RTG or heater is used to heat the propellant storage tank. Gas jets provide low thrust (<10 Newtons) and low total impulse (< 4,000 Newton Seconds). Cold gas jets provide specific impulses in the range of approximately 65 to 75 seconds and warm jets provide specific impulses of about 105 to 230 seconds [Sutton 1986]. Gases such as hydrogen can provide significantly higher specific impulses, but in general any advantage gained is negated by the bulk of the storage systems necessary to contain them due to their very low densities. Table 8.3.3 contains some possible gases for use in gas jets.
Hybrid engines
Hybrids combine some of the advantages of liquid and solid propellants. The liquid (usually the oxidizer) is stored in one container with the solid (usually the fuel) in a second. The separation of propellants eliminates the dependence of burning time on the grain area while the absence of oxidizer in the solid grain improves its structural properties. The hybrid combines the start-stop advantages of liquid propellants with the high density, instant readiness, and potentially high thrust of solid propellants. The drawback is designing the proper flame pattern without degrading performance. The theoretical specific impulse of a solid lithium/liquid flourine and oxygen mixture engine is about 375 seconds.
NON-CHEMICAL SYSTEMS
Non-chemical systems function fundamentally the same as chemical systems, but the difference is in the source of energy for the working fluid and in the working fluid itself. Non-chemical systems obtain their energy by processes other than the combustion or decomposition of chemical reactants. Non-chemical systems energy sources include such things as thermal, electrical, magnetic, electromagnetic, and other forms of energy. The working fluid is also different. The working fluid can be a gas such as hydrogen, or it can even be ions. In the following sections, various non-chemical systems will be discussed. The first system to be examined is the nuclear thermal rocket.
Nuclear Thermal Rockets
Nuclear thermal rockets obtain their energy in the form of thermal energy from a nuclear reaction. Thrust is produced by feeding a gas over a nuclear reactor and then accelerating this heated gas through a converging-diverging nozzle, just as the reaction products are in a chemical system. The specific impulse increases with higher chamber temperatures and with a decrease in the molecular weight of the exhaust gas, therefore hydrogen is usually used as the working fluid, or propellant. A solid core nuclear rocket can produce specific impulses up to 1000 seconds, which is twice that of the best chemical rockets. Several solid core nuclear rockets were developed and tested during the 1960's. A second type of nuclear rocket is the gaseous core rocket. This type engine operates at much higher temperatures and produces Isp from 1,000s to 6,000s (Angelo). These rockets have not been tested, but do look promising for interplanetary missions. The main advantages of theses systems are the very high specific impulses of which they are capable coupled with
the fact that they produce high thrust and are reusable. On the other hand these systems use nuclear reactors and thus require much shielding, especially for manned missions, thus negating some of the advantages. There are also strong political and environmental groups which are opposed to the use of nuclear energy for any purpose.
Electric Thrusters
Electric thrusters' source of power is electricity. This electricity may be used to supply thermal energy or electrostatic/electromagnetic energy to the working fluid. The following two sections discuss both electrostatic, electrothermal, and electromagnetic systems.
Electrostatic/Ion
The working fluid in ion engines are ions. Ion engines develop very high specific impulses on the order of 2,000 to 10,000 seconds but at very low thrust levels. These engines produce very low accelerations over long periods of time. Ion engines produce thrust by pumping a neutral propellant into an ion production chamber where ions and electrons are separated into two different streams. The ions pass through a strong electrostatic field and are accelerated into an exhaust stream. The thrust is the total reaction to the accelerating forces [Hill]. A large amount of electricity is required to produce this electrostatic field, therefore a nuclear reactor is often used to supply the power to the engines. Another option is to use solar power to generate the electricity; this is given the name, "Solar electric ion propulsion." Figure 8.4.1 shows a schematic of an ion engine [Sutton 1986].

Electrothermal
Electrothermal motors operate by adding thermal energy directly to the propellant and expanding it through a conventional nozzle just as in chemical and nuclear thermal systems. A simple example of such a motor is the resisto jet. In this example electricity is used to resistively heat filaments which in turn heat the propellant. The thermal limitations of the filaments allow such a design to reach a maximum specific impulse of about 800 seconds. Another more common form of electrothermal motor is the arc jet thruster. In this design the electrical current is made to pass directly through the propellant to heat it. This bypasses the limitations of the heater filament and theoretically allows specific impulses of approximately 1500 seconds to be achieved. Unfortunately problems arise due to the fact that the electricity is passing through the propellant thus causing it to ionize. This dissociation of the propellant in turn leads to other problems.


Electromagnetic (Magnetoplasmadynamic)
Electromagnetic thrusters were developed to try to circumvent the ionization problems of the electrothermal motors. In this concept the Lorentz force (XB) resulting from the interaction of an electrical current with a magnetic field is used to add to the electrical energy directly in kinetic form. Significant engineering limitations have limited the advantages and potential of this system. These systems can provide high specific impulses, but at the expense of high power requirements, low thrust levels, and thus long travel times
EXOTIC/THEORETICAL SYSTEMS
Included in this category are systems which are novel or in the distant future. Some of the systems include in this system use solar wind or thermal radiation as the means of propulsion. Another uses discrete solid particles as opposes to a fluid for the propellant. The first system to be discussed is solar sails.
Solar Sails
The solar sail is a unique propulsion system that takes advantage of the Sun's radiation pressure. The Sun produces a gradient of radiation pressure, with the magnitude decreasing outward from the Sun. This pressure produces a small force on an object in space, but it is usually regarded as nothing more than a perturbation to the object's orbit. However, a large sail (sometimes on the order of several square kilometers in area) could exploit this free solar wind. For example, the sail could provide a thrusting force for a ship on an Earth-Mars trajectory. Although the thrusting force is still small, it can now be controlled and made to contribute to the mission. A great advantage of the solar sail is obviously that it doesn't require any propellant, but there are several drawbacks as well. The sail is virtually useless beyond the orbit of Mars, as the radiation pressure dwindles at large distances from the Sun. Also, the sheer size of the solar sail will pose new problems in construction and control. Finally, a large object such as a sail is likely to collect debris and perhaps suffer damage from such encounters. A small scale solar sail experiment is slated for a Space Shuttle mission in the early 1990's. Solar sails potentially can provide very large specific impulses and are reusable, but they produce very small thrust levels and the structures are quite fragile and massive.
Solar/Laser Thermal Propulsion
In a solar or laser thermal rocket, solar or laser light is collected and focused to heat a propellant working fluid such as hydrogen. The collector mirrors are very large structures which serve to concentrate the light energy on the propellant. Lasers could be based on the ground, in an airplane or in space. Very high pointing accuracy is required for the laser, and a laser thermal rocket could be used only near the source of the laser. Specific impulses in the range of 800 to 1,500 seconds are possible and thrust to weight ratios of about 10-2 are expected.
Rail Guns
Rail guns use electromagnetic fields to rapidly accelerate payloads to very high velocities on the order of kilometers per second and could provide specific impulses in range from 600 and 1,500 seconds. Rail guns require very large and complex systems to set up the electromagnetic fields and manipulate them to produce the high accelerations, thus suggesting probably a stationary propulsion system. But on the negative side, payloads would be subjected to very high g-loadings and large thermal loads due to atmospheric heating if used from the earth. Environmentalist would also be quite displeased with the multiple sonic booms which would be created.
Antimatter
Antimatter propulsion is a theoretical concept in which propulsion is derived from the annihilation energy of matter-antimatter reactions. The theoretical specific impulse that could be obtained from such a reaction is on the order of 3.0 x 107 seconds and velocities approaching the speed of light could be obtained. The major obstacles in the production of antimatter engines are in the production and storage of the antimatter. Current production rates would require 1.0 x 108 years to produce one kilogram of antiprotons. Magnetic storage would also be required to contain the antimatter and prevent contact with any physical container.
SELECTING A PROPULSION SYSTEM TYPE AND SIZE
This section provides a spacecraft designer with information for obtaining a first approximation of the engine size and engine type that will best suit the mission requirements. Figure 6.1 has been provided to illustrate some of the important characteristics of different propulsion types. It may be used to eliminate one or more propulsion types from consideration, narrowing the scope of the search for the best propulsion system type based only on top-level mission characteristics. The following steps will help determine the mass of the engine and the propellant mass that will be required for the mission:
1. Calculate the change in velocity (DV) that the propulsion system will be expected to deliver.
2. Estimate the maximum thrust the spacecraft structure can withstand.
3. Refer to the selection tree presented earlier to select the type of engine that may be best suited for the spacecraft based on performance requirements and structural strength.
4. Choose a specific impulse which falls within the range of the engine selected.
5. Use the graph of Thrust vs. Engine Mass shown in Figure 2 to obtain an estimate of engine mass.
6. Use the following equation to estimate the propellant mass for one burn.
7. Calculate the total impulse.
8. Consult the Propulsion System Database to select a specific engine that meets the criteria of specific impulse, total impulse, and/or thrust that you have determined.
a) Solid Motors
(1) Eliminate all solid motors that produce more thrust than the spacecraft structure will be designed to withstand.
(2) Pick a motor with the same or slightly more total impulse as found in equation (2).
(3) Take the motor mass, propellant mass, and specific impulse from the Propulsion System Database.
(4) Take or recalculate the total impulse using the engine mass and make sure it is high enough to accomplish the mission. If not, select another motor.
b) Liquid, Nuclear Thermal, Ion
(1) Eliminate all engines with too much thrust.
(2) Choose an engine with an acceptable thrust.
(3) Divide the total impulse from equation (2) by the thrust of the engine to get the time of burn.
(4) Multiply the time of burn by the mass flow rate the engine to get the mass of the propellant.
(5) Check to see if the mission DV's can be obtained using the following equation:
Figure 6.2 shows the relationships between engine thrust and engine mass for three different types of liquid propellants, storable, LO2 Storable and LO2/LH2. These relationships can be used to estimate the masses of engines with thrusts different from those in the PSD.
SPACE SYSTEMS DESIGN CLASS PROPULSION DATABASE
The Propulsion System Database (PSD) is a personal computer based database of specific propulsion systems. It was designed to assist spacecraft designers in selecting a set of suitable propulsion system candidates. It contains specific information concerning mass and dimensions for already developed or under final development engines . The PSD remains to be completed . The main source of information is a catalog displaying characteristics of most of the current engines [Wilson 1991].
Database Characteristics
There are three distinct databases in the PSD : liquid propellant engines, solid propellant engines, and monopropellant thruster. "Exotic" means of propulsion are not yet taken into acoount. For some of the engines, only a few data where available. In general, when for a field several values are given, the value chosen is the biggest one in order to have a conservative estimation for the sizing of the spacecraft design (for example, the gas pressure in the chamber). The PSD has only to be used for a primary evaluation and choice for the design of the propulsion system. The manufacturer should be asked for further information, including new improvements of the current engines. Most of the engines described exists in many different versions.
Field and description
Name (alphanumeric) - This field contains the common name or trade name of the particular engine or engine type.
Manufacturer (alphanumeric) - Engines that have been designed or produced by a particular company have the name of the manufacturer listed in this field.
Use (alphanumeric) - A primary use of the engine will be listed, when available.
First flight : year of the first launch using the oldest version of this engine
Dry mass (Kg) : for bi- and mono- liquid propellant engines only
Total mass and propellant mass (Kg) : for solid propellant engines only
Length (mm)
Diameter (mm)
Mounting : the way the engine is fixed on the structure.
Engine cycle : the feeding of the engine
Oxidizer, Ox. flow, Fuel , Fuel flow and O/F ratio : for liquid engines only
Propellant : for solid prop. engines and thrusters only
Propellant flow : for the thrusters only.
Thrust range (N)
Nominal Thrust (N)
Specific Impulse (Isp) at sea level and in vacuum (s)
Expansion ratio
Pressure in the combustion chamber (atm)
burn time (s) : for restartable engines, this is the total burn time possible over
several starts. A very high burn time means that is unlimited : that the case
for most of the small thrusters.
ANTIPROTON-CATALYZED MICROFISSION/FUSION
In 1992 our group observed large fission and neutron yields from antiproton annihilation at rest in a natural uranium target. Calculations indicate that short bursts of antiprotons could induce temperatures of several keV in a small compressed pellet. These conditions are appropriate for ignition of a hydrogen fusion burn within the microsphere. Targets with yields up to 302 GJ have been considered, with compression provided by light ion beams or lasers. Baseline parameters for ignition are: antiproton energy, 1.2 MeV; number, 1011; pulse length, 2 ns; and deposition volume, 1 mm. An experiment at the Phillips Laboratory to demonstrate subcritical antiproton-catalyzed microfission is in progress
Most of the energy from the microfission/fusion process is in the form of radiation and hot (35 keV average temperature) plasma. Energy is produced in a target consisting of about 3.0 g of nuclear fuel. The nuclear fuel is in a molar ratio of 9:1 of DT:U(235). Initially, the proportions of energy produced in the target are 83% radiation, 15% neutron kinetic energy, and 2% random ion and electron kinetic energy. Since most of the energy is in the form of high energy radiation, a high Z material (WLS) is desired to absorb and reradiate at lower frequencies (temperatures). The purpose of this is to optimize ablation of a thruster shell exposed to this radiation. The WLS is a spherical shell of 200 g of lead, which has a K-shell absorption edge near the peak of 115 keV for a spectrum with an average temperature of 35 keV. Of the 302 GJ of energy generated in the target, 247 GJ is absorbed by the WLS. This energy is distributed over the WLS volume according to a stellar photosphere model,8 initially 5.6 keV at the center and 2.3 keV at the surface. Since only a thin skin on the surface of the shell radiates, most of the lead is near the 5.6 KeV temperature, which corresponds to an ionization level of Z*=75. This temperature is not high enough to remove inner shell electrons, thus enabling continuing K-shell absorption of radiation. Energy distributions of photons radiated from the photosphere around the WLS show a very significant shift of radiation down to a mean value of about 1 keV energy (204 GJ), which greatly enhances the absorption of this radiation in the ablative thrust shell used in the ICAN-II spacecraft.
ICAN-II PROPULSION SYSTEM
The basic premise of the explosion concept is that a spacecraft equipped with an adequate shock absorbing apparatus and a means of safely intercepting debris from a nuclear explosion could use the explosion energy to propel the craft at both high thrust and high Isp. Historically, the first serious effort at this type of propulsion was the ORION pusher-plate system. A more sophisticated system employing a polyethlyene canopy tethered to a winch (MEDUSA) has been proposed by J. Solem.

In the current version of the ICAN-II concept, a sector of a spherical silicon carbide (SiC) shell of 4 m radius is used to intercept radiation from the explosion. This radiation heats the inner surface of the shell to keV temperatures, and the resultant expanding plasma produces thrust. A schematic version of the ICAN-II spacecraft, including the engine at the aft end, is shown in Figure 1. Estimates of component masses for a 120 day, V = 100 km/s Mars mission (RT) are given in Table 1.


OTHER FEATURES OF ICAN II
Additional major systems and performance characteristics of ICAN-II are reviewed below:
Radiation Control and Power Systems
Figure 2 shows a schematic diagram of the ICAN-II radiation and power systems. Radiation damage to the ICAN-II vehicle can result from neutrons, which pose a threat to the stored targets, antiprotons, and the crew of the vehicle. In order to prevent this, 1.2 meters of lithium hydride shielding (Power Shield) is required. In addition, 2.2 meters of shielding Crew Shield) is needed to limit crew exposure to ~ 30 rem over the duration of a mission. Finally, part of the intense blast of neutrons from the pellet ignition is absorbed by the Power Shield to drive a 10 MW electric generator, which provides power for the ion drivers and other systems on the spacecraft. A liquid droplet radiator Figure 3) has been designed for expelling the excess 60 MW of heat from the Power Shield into space.
Thrust and Isp
Figure 4 shows the thrust and Isp for a 1 Hz firing rate. For a V of 100 km/sec and an Isp of 13,500 seconds (200 g WLS), 362 metric tonnes of propellant are required for a 345 metric tonne ICAN II dry mass (see Table 1). With a 200 g WLS, the thrust is about 100 kN, which accelerates the outbound craft to a 25 km/sec V in 3 days. For 800 g of ejected mass, about 30 ng of antiprotons are required. Hence, ICAN-II could be fueled with one year's production of antiprotons at Fermilab, estimated by ourselves to be approximately 140 ng by the year 2000.
Energy Utilization
It is interesting to investigate how energy from the exploding pellet is productively utilized in ICAN-II, or more generally in a fusion-driven spacecraft with ablative thrusters. The results are given in Table 2, where 7.1 GJ (7.1 GW @ 1 Hz) energy is
imparted to actual forward kinetic energy of the spacecraft.

.
. EXAMPLES OF INTERPLANETARY AND EXTRAPLANETARY MISSIONS
Utilizing vehicle performance parameters presented above, three potential ICAN-II missions were analyzed. As an intermediate step to a full non-impulsive analysis, simulations of vehicle trajectories within planetary gravitational spheres of influence were performed by modeling vehicle thrust and solar gravity as perturbations. The results indicate that the majority of the V was gained within the planetary spheres of influence, permitting the design of interplanetary trajectories using impulsive maneuvers at the endpoints. Missions to Mars, Jupiter and Pluto were investigated, and the results are presented in Table 3. The short transfer times significantly alleviate psychological and physical dangers to the crew. A total V requirement of 120 km/s was stipulated to provide a large launch window every two years, although the mission can be completed with as little as 70 km/s if departure is timed correctly.


PRODUCTION OF ANTIPROTONS
Antiproton sources exist worldwide at two sources, CERN in Geneva, Switzerland and Fermilab, in Batavia, Illinois. These two laboratories utilize high energy proton synchrotron accelerators, with accumulator storage rings attached to collect antiprotons produced by collisions of protons on targets. Presently, Fermilab collects 6 x 1010 antiprotons per hour in its Accumulator. This means that in one year of dedicated production, it could produce a maximum of 0.85 ng of antiprotons. A new and funded facility, called the Main Injector, will turn on in 1999, with a maximum annual production capacity of 14 ng. A new Recycler Ring presently under construction and located inside the Main Injector ring will increase the collection rate by another factor of 10. This would place Fermilab in the 100 ng range, making it attractive for future space applications.
TRAPPING ANTIPROTONS
Antiproton trapping work has been done at the Low Energy Antiproton Ring (LEAR) at CERN.12 LEAR has provided low energy antiproton beams, presently not available at Fermilab. The 5.9 MeV antiproton beam is degraded down to an energy of 10-30 keV and injected into a large Penning trap. The antiprotons are trapped radially by the magnetic field, and Figure 5. Sensitivity of 120-m Radiotelescope on ICAN-II to Signals from the Center of the Galaxy. axially by the two confining electrostatic potentials. The harmonic frequencies of these two motions are about 50 and 5 MHz respectively. A third harmonic "magnetron" motion is also present. This precession around the direction of the E_ x B_ vector is at a rate of about 80 kHz. Measurement of the actual number of antiprotons trapped is done by lowering the potential of the far electrode, allowing the antiprotons to spill out of the trap and annihilate into charged pions. Observed linear correlations between the number of pion counts and the number of antiprotons injected into the trap show that at least 106 antiprotons per injection shot from LEAR have been trapped. Electron cooling then permits collection of successive shots from LEAR, for example 10 successive shots would yield 107 antiprotons in the trap. Electron cooling is done by injecting electrons into the trap, where by collisions they absorb energy from the antiprotons. This energy is released by the electrons as they spin around the magnetic field in the form of synchrotron radiation. The data demonstrate lifetimes of up to several hours, corresponding to vacua of less than 10-11 Torr. TRANSPORTING ANTIPROTONS TO SPACE
For space propulsion applications, 140 ng of antiprotons corresponds to about 1017 antiprotons. One possible scenario therefore would be to transport 103 traps into space, each holding 1014 antiprotons. It is likely that these 103 traps would be integrated into a common cryogenic system. Scale-up from traps holding 107 antiprotons to 1014 antiprotons will not be trivial. Traps presently in use have a Brillouin limit of about 1011 antiprotons/cc. Therefore, a trap with a volume of 1 liter can hold the required number of antiprotons. We are presently building a portable antiproton trap.13 It is designed to carry up to 109 antiprotons for 4-10 days. A schematic drawing of the trap is shown in Figure 6. It is a prototype for a trap capable of carrying 1014 antiprotons for up to 120 days (duration of a round trip mission to Mars). The portable trap is one meter tall, 30 cm across, and weighs 125 kg. It operates at 4K temperature, supported by cryogenic nitrogen and helium reservoirs, and has a unique feature that the confining magnet is made of permanently magnetic SmCo materials, which should prove to be robust. Test results to date are very encouraging. Up to 40 million electrons have been trapped for sixteen hours. H2 gas (~ 1 mole) has been injected and the electron gun turned on. Bombardment of the gas by the electrons produces various charged ion species, including small numbers of H+ ions. The storage lifetime has been measured by extraction into a channeltron detector (Fig. 7). Lifetimes of up to 103 seconds have been observed. The electron and H+ lifetime results are consistent with a vacuum in the inner trap of 10-10 Torr. We are currently installing a high current H+ injector to permit storage of 106 ions. When tests of the system operating on a stand-alone battery system are completed, a portability test of H+ ions will be made. We estimate that vacuums presently achieved correspond to an antiproton lifetime of 14 hours. We foresee improvements of vacuum by a factor of five, arriving at a 4 day storage time in the near term.
Instabilities set in when the charged antiproton Coulomb energy density exceeds the magnetic (Penning trap) energy densities. Since there are practical limits to fields that can be supported, the next step is to prepare accumulations of large numbers of antiprotons in the form of electrically neutral atoms, such as atomic antihydrogen. Within the next two years these atoms will be synthesized at CERN by injecting positronium atoms, bound electron-positron pairs, into a trap filled with antiprotons. Initially we hope to form and confine thousands of antihydrogen atoms in a Pritchard-Ioffe trap, consisting of a vacuum cylinder within a quadrupole magnet, augmented with confining pinch coils at each end. Confinement is provided by the interaction of the atomic magnetic moment with the inhomogeneous magnetic field. This technology is currently available from laboratories studying atomic hydrogen where densities of >1014 atoms/cc have been achieved. Although these densities are much higher than allowed by Penning traps, instabilities exist which prohibit their use at high densities for long term accumulation. The next step therefore involves forming condensates of electrically neutral molecular antihydrogen, either in liquid or solid form, which would provide densities approaching 1023 atoms/cc; 140 ng of antihydrogen would constitute a spherical volume of about 60 micrometers radius.
Confinement of antimatter has been the subject of extensive studies.14 Since liquid or solid antihydrogen is diamagnetic, levitation within a confining vessel could be provided by a magnet of modest size.15 Serious technical issues include annihilation of surface atoms with residual gas in the confining vessel, and sublimation of surface atoms with resultant annihilation on the walls of the confining vessel. In the latter case, the annihilation could eject matter from the walls, which in turn annihilates with the antihydrogen, starting a chain reaction.16-17
A photon sail is a spacecraft accelerated by the momentum of the electromagnetic photons that are reflected from it. Types of photon sailcraft are laser sails (which are pushed by laser beams), maser sails (which are pushed by collimated microwave beams) and solar sails. The solar sail is accelerated by momentum transfer from photons emitted by the Sun that strike the sail. If the sail is fully opaque (non transmissive), the radiation pressure of the solar photons impinging against a sail oriented with its reflective surface normal to the sunlight can be written:
Rad. Pres. = (1 + Ref sail) SF / c Newton / m2 , (1) where is Ref sail is sail reflectivity, SF is the solar flux striking the sail (1,368 watts / m2 at the Earth’s distance from the Sun—the Solar Constant S.C.) , and c is the speed of light (3 x 108 m/sec). The solar flux striking the sail surface varies with the square of the inverse of the distance to the Sun’s center.
Sail Kinematics
If one multiplies Eq. (1) by the sail area normal to the Sun, Asail in square meters, the resulting equation yields the force of the sunlight on the sail surface. Applying Newton’s Second Law and dividing this result by spacecraft mass Ms/c kilograms, the sailcraft’s acceleration due to solar radiation pressure is obtained :
ACCs/c = (1 + Ref sail) SF Asail / (Ms/c c) = (1 + Ref sail) SF / (s/c c) m/s2 , (2) where s/c is the sailcraft’s areal mass thickness in kilograms per square meter. A convenient Figure of Merit for sail designers is the sailcraft Lightness Factor, LFs/c, which is the ratio of solar radiation pressure acceleration to solar gravitational acceleration on the sailcraft. From Newton’s equation for Universal Gravitation, LFs/c = (1 + Ref sail) SF Rsun 2 / (G Msun s/c c), (3) where G is the Gravitational Constant 6.67 X 10-11 N m2 kg-2 , Rsun is the distance of the sailcraft from the sun’s center (1.5 X 1011 m at the solar orbit of the Earth—1 Astronomical Unit or 1 AU) and Msun is the Sun’s mass (1.99 X 1030 kg). Substituting in Eq. (3), we obtain the simplified result :
LFs/c = 0.000773 (1 + Ref sail) / s/c . (4) At the Earth’s distance from the Sun, a fully unfurled sail with a Lightness Factor of 1 is accelerated by solar radiation pressure at 5.92 X 10-3 m/s2 or about 6 x 10-4 Earth surface gravities.The areal mass thickness of such a sailcraft will be 0.00146 kg / m2 . Solar sails must therefore be very light and very reflective. If the sailcraft described in the previous paragraph has a mass of 500 kg, the sail area normal to the Sun must be about 3.42 x 105 square meters. This sail will be considerably larger than a football field!
What if the Sail is Not Normal to the Sun?
Let’s call the angle between incident sunlight and the normal to the sail “”. If the sail is normal to the sunlight, = 0 degrees. If the sail is not normal to the Sun, the normal value of solar irradiance (in watts / m2) striking the sail is multiplied by cos , according to Lambert’s Law. Since the crosssectional sail area normal to the incident sunlight is also reduced by cos , the sailcraft acceleration at Sun-sail angle (ACCs/c,), is related to the sailcraft acceleration for the case of perpendicular sunlight (ACCs/c,) by the equation:
ACCs/c,= ACCs/c,cos2
When the sail is not normal to the Sun, there will be two components to the sailcraft’s solar-radiation-pressure acceleration vector. These are ACCs/c, rad , the component of spacecraft acceleration radial to the Sun and ACCs/c, tan, the acceleration component tangential to the spacecraft’s orbit around the Sun. Mathematically,
ACCs/c, rad = ACCs/c,cos , (6a)
and
ACCs/c, tan = ACCs/c,sin . (6b)
In the 2-sail solar photon thruster discussed below, Lambert’s Law does not apply because the main, collector sail is always normal to incident sunlight and all sunlight striking the main collector is directed against the thruster, regardless of thruster angular orientation. For such a sailcraft, the cos2term in Eq. (5) is replaced by cos .
Sail Thermal Effects
As well as being low in mass and highly reflective, the sail must be constructed of heat-tolerant material. This is because all of the solar energy absorbed by the sail must be radiated from the sail as infrared electromagnetic radiation. The radiant power absorbed by an opaque sail oriented normal to the Sun is written:
Pabs = (1-Refsail) SF Asail watt , (7)
Since absorbed electromagnetic radiation can be reemitted as infrared from both sides of
the sail, the sail’s radiant emittance can be expressed as :
Wsail = (1-Refsail) SF / 2 watt / m2 . (8)
The Stefan-Boltzmann Law for greybodies can be used to relate sail radiant emittance to sail absolute radiation temperature Tsail :
Wsail = Tsail
4 watt / m2 , (9)
where is the Stefan-Boltzmann Constant (5.67 X 10-8 W m-2 K-4 and is the sail material’s emissivity. (For a blackbody, = 1). Equating Eqs. (8) and (9) and substituting SF = 1,368 watt / m2 at 1 AU, the Earth’s average distance from the Sun, Tsail = 331 [(1-Refsail) / ] ¼ degrees Kelvin. (10) The maximum sail radiation temperature must never exceed the melting point of the sail material. For most materials, the maximum allowable radiation temperature is a few hundred degeees Kelvin less than the sail-material melting point.
Some Sail Configurations
Figure 1 shows six proposed sail configurations. A disc sail consists of a circular sail film supported and stabilized by a series of spars as shown. The payload is often at the center of the spar structure. Square sails are square or rectangular (when viewed from top or bottom). The spars and payload are often arranged in a manner similar to disc sail designs. Parachute sails physically separate payload from sail. These are attached by a series of cables, as shown. The parabolic sail or solar-photon thruster is a multi-sail configuration. A main collector normal to the Sun focuses sunlight upon a smaller thruster. The thruster can be steered allowing the spacecraft to alter the direction of reflected sunlight and thus be more maneuverable than other sail configurations. Although many sail designs rotate so that centripetal force can aid in sail unfurlment, the heliogyro spins like a gyroscope, although more slowly. In some designs, sail film rolls out from canisters near the center along spars within the sail blades. Hollow body or pillow sails are inflatable devices with the surface facing the Sun coated with a reflective layer. Although these may be easier to unfurl than other sail concepts, they have the disadvantage of being more massive.
Solar-Sail Limitations and How to Overcome Them
One major limitation to solar-sail application in the outer solar system and beyond is the inverse-square-law of solar irradiation. As the sail doubles its distance from the Sun, the solar flux and radiation-pressure acceleration fall by a factor of 4X.
In concept, this effect can be compensated for by replacing sunlight in the outer system and beyond by collimated radiation. Studies have indicated that solar-powered visible and infrared lasers or collimated microwave masers in the inner solar system can project beams against sails very far from the Sun. Such sailcraft should properly be called
“light sails” rather than solar sails.
Solar Sail Maneuvers and Orientations
Solar sail maneuvers can be classified using analogues with maneuvers performed by conventional oceanic sailing craft. As shown in Fig. 2, a sailcraft can be oriented with its sail normal to the incident sunlight. In such an orientation, reflected solar photons produce maximum thrust on the sail surface. The craft “runs with the (solar) breeze”.
If the sail is oriented at an angle to the incident sunlight, the sailcraft’s thrust willbe at an angle to incident sunlight. In this orientation, the spacecraft “tacks”. In a two-sail solar photon thruster, thrust can be greatly reduced by reflecting sunlight from the thruster back towards the main collector sail. This orientation is analogous to the “hove-to” maneuver performed by an oceanic sailcraft to slow the craft under high-wind conditions.
Solar-Sail Materials
A number of choices exist for solar sail materials. Many of these are reviewed in the references. For an Earth-launched near-term sailcraft, one approach is to create the sail using a sandwich of three materials. A metallic, highly-relective (front) layer faces the Sun, followed by a layer of flexible, temperature-resistant plastic. The back layer is composed of a highly emissive material such as chromium. The function of the layer facing the Sun is to reflect as much of the sunlight as possible (up to about 90%). The plastic layer’s function is to improve sail flexibility during the sail unfurlment process. Sunlight absorbed by the front layer is reemitted as infrared electromagnetic radiation from the back emissive layer. Some current three-layer sail designs have areal mass thicknesses less than about 10 grams per square meter and are capable of operating within
0.1 AU from the Sun. The sail thickness is typically of the order of a few microns. Additional sail strength can be achieved by including strips of metallic ribbing within the sail structure. These might minimize the effects of micrometeoroid impacts. Some researchers have examined the possibility of using plastics that would rapidly degrade when exposed to solar ultraviolet radiation. In a sail constructed of this material, the “sandwich” layers would be reflective, emissive, and finally UV-sensitive plastic. This approach could greatly reduce sail areal mass thickness. The thinnest Earth-launched sails might consist of a stack of strong, temperatureresistant composite fibers. Perhaps coated with a reflective layer, the areal mass thickness of such a sail could be less than 1 gram per square meter. A sailcraft constructed using this material could operate well within 0.1 AU. Sail unfurlment may be an issue for very thin Earth-launched solar sails. Ultimate solar-sail performance requires a space-manufacturing infrastructure. Using vapor-phase deposition, large-metallic sheets 20-30 nanometers thick could be produced in space. Areal mass thickness of 0.05 grams per square meter might be possible. Some metallic hyperthin sails could (theoretically) approach the Sun within the 0.05 AU of the Sun’s center. Nanotechology might, in the farther future, allow for the creation of perforated or
mesh solar sails. If the perforations are substantially smaller than the wavelength of incident light, low mass and high reflectance might combine to greatly increase sail performance.
Solar-Sail Missions
The solar sail is unique among currently-feasible advanced propulsion schemes because it requires no propellant and can accelerate endlessly without thrusting under the influence of sunlight or some other source of electromagnetic radiation. Once unfurled, a solar-sail constructed of material that does not readily degrade in the space environment, should have a very long operational life. Many space missions are possible using solar sails that are more difficult or impossible using other in-space propulsion modes. In general (unless power-beaming is applied), these are either inner solar-system missions or missions in which spacecraft acceleration is accomplished within the inner solar system. The solar sail could accelerate to velocities of 100 km/sec or higher after unfurlment during a close solar flyby. But (since the solar flux will be reduced from its near-Earth value by a factor of 25X at the 5-AU solar distance of Jupiter) the solar-sail will be ineffective for stopping at outer-solar-system destinations. In low-Earth orbit, atmospheric drag is also an issue for the high-area, low-mass solar sail. Even solar-photon thrusters may be drag limited to unfurlment orbits in excess of 500 km above Earth’s surface.
Near-term scientific missions
In the near-term, solar-sails in the 6-10 grams per square meter range, with areas of about 104 square meters are under consideration for inner-solar-system solar observer missions. Stationed for years between the Earth and Sun, such craft could give early warning of solar flares approaching the Earth. If they are constructed of material capable of withstanding radiation produced by repeated passes through Earth’s Van Allen radiation belts, sail configurations such as the solar-photon thruster could maintain constellations of satellites at various locations within Earth’s magnetosphere. The payloads of these craft could monitor the near-Earth plasma environment for examine space-magnetic interaction s with the terrestrial environment. The solar-photon-thruster sail configuration could enable the “pole sitter,” a spacecraft situated at lunar distances over or nearly over one of Earth’s poles. Pole sitters have application to high-latitude communication, climate studies and Earth viewing. Sails could be considered for payload delivery or round-trip visits to inner-solar system planets as far out from the Sun as Mars or for launching probes to the outer solar system. Fast flybys of outer-solar-system objects such as Kuiper Belt cometoids (30-50 AU from the Sun) are possible. Perhaps the most exciting near-term possibility is a sailcraft equipped with a sail that is unfurled within the orbit of Venus or perhaps Mercury and then accelerated out of the solar system on a trajectory capable of reaching the heliopause (the limit of the Sun’s influence--at 250 AU from the Sun) within a few decades. Such an interstellar probe could monitor in situ the interaction between the Sun and the galaxy.
Space-commercialization and Earth-protection missions
As humanity’s space infrastructure matures, Earth-launched or spacemanufactured solar-sails will offer opportunities to those seeking to develop the spaceenvironment or to protect the Earth from cosmic collisions. Near-Earth asteroids and comets could be explored and mined using solar sails. If advanced terrestrial and space telescopes can provide decades-long warnings of collisions between Earth and near-Earth objects (NEOs), solar sails-could be unfurled around or near the offending asteroids or cometary nuclei. Radiation pressure on sails attached to NEOs could alter collision trajectories to near-miss trajectories. Alternatively, sails could be used as concentrators to heat NEO surfaces enough for trajectory-altering jets of material to be ejected.
To the Oort Belt and the stars!
As 0.1 gram / square-meter sails capable of withstanding very close perihelion passes are developed in the farther future, solar-system exit velocities in excess of 300 km/sec will become possible. Then, space-mission planners could design robotic missions to nearer Oort Belt comets within a few thousand Astronomical Units of the Sun with flight times of a few decades. Ultimate space-manufactured solar sails will allow the possibility of Earth-Mars roundtrips with durations measured in months rather than years. The same technology may allow visits to the nearer stars at solar-system exit velocities higher than 1,000 km /sec. The one-way travel time to the Sun’s nearest interstellar neighbors (proxima and Alpha Centauri) will be about 1,000 years for such spacecraft. Advances in laser and maser power-beaming technologies may eventually substantially reduce interstellarvoyage durations.