Space propulsion has changed little since mankind took its first tentative steps into space. Even with the incremental advances in the efficiency of chemical fuels; the basic nature of rocketry is still defined by the basic Delta V Rocket Equation with all its limitations; be it the powerful boosters used to obtain orbital velocity or the low impulse Ion Thrusters used to power deep space missions. This Neanderthal approach to propulsion limits both the potential flight parameters of deep space missions and the life span of earth orbiting satellites.
In earth orbiting satellites, the electronics of the satellite may last indefinitely; but the useable lifespan of that satellite is limited by the availability of on-board propellants used for orbital maintenance. Once the chemical propellant is exhausted, the satellite no longer has the capability of maintaining proper station.
The International Space Station is dependent upon chemical propellants to offset orbital decay. The need and use of these chemical propellants increases the potential for catastrophic accident, increases the cost of operational maintenance, and requires the commitment of launch capacity for that purpose.
Interplanetary and deep space missions face similar limitations inherent to dependence on chemical propellants. Although gravitational assist has been a regular tool used in both navigation and imparting changes in specific orbital energy; obtainable velocities, launch windows, and other flight parameters remain severely limited by dependence upon the same Newtonian Propulsion methods used by the ancient Chinese to power their rudimentary rockets. Even Ion Propulsion, which uses electro-magnetic acceleration of the ion fuel to achieve impulse, is still a type of Newtonian Propulsion where the total energy imparted is limited by exhaust velocity and total available fuel mass as defined by the basic rocket equation. Newtonian Propulsion may have gotten us to earth orbit and beyond; but it will be Electro-magnetic propulsion that will carry us to the stars. In the mean time, its development will allow us to achieve flight parameters unimaginable when considering only chemical propellants.
The Ampere Defined As Magnetic Force:
Prior to 1948 the ampere was defined, based on Faraday’s Law Of Electrolysis, as the amount of unvarying current, that when passing through a solution of silver nitrate, deposits silver at a rate of .00111800 grams per second.
The ampere was redefined in 1948 as the amount of unvarying current, that when being carried by two infinitely long conductors separated by one meter, would generate a magnetic force between the conductors of 2 X 10-7 Newton per meter of length. This is the Standard International definition of an Ampere.
Electro-magnetic Propulsion and it’s inverse, Electro-dynamic Braking, when combined with the now and near term future technologies related to super conductivity, dielectric capacitance, and other related technologies; will introduce a new paradigm in space propulsion.
The Fundamentals Described As A Space Based Experiment:
A simple space based experiment to demonstrate the basic principles of electro-magnetic propulsion is easily imagined. In this experiment a simple coil, a number of accelerometers, a polarity reversing switch, power source, and radio telemetry is used to determine the earth’s electro-magnetic field strength at the range of the experimental package. It would be most advantageous if the coil length is as great as possible. The coil is circuited in series with the polarity reversing switch and the power source. The accelerometers serve to activate the polarity reversing switch. The experiment is then suitably packaged and conventionally launched to a low inclination orbit. The experimental package is then positioned so that the field coil of the package is aligned so the coil will be at maximum repulsion with earth’s electro-magnetic field when the coil circuit is initially energized.
So positioned, when the switch is initially closed and power is applied to the coil there will be two vector forces acting on the coil.
Since the coil is aligned in repulsion with earth field, one force will be acting along a line that is perpendicular to the earth’s North/South polarity (approximating the line of orbital radius) and will translate to an acceleration vectored along the orbital radius converting circuit energy to increased gravitational potential.
The other force acting on the field coil will translate to torque causing the field coil to begin to rotate about the central axis of the coil length as it begins to align towards the magnetic equilibrium position relative to earth field; that being one of maximum attraction and 180 degrees relative to the maximum repulsive position. In doing so, some of the electrical energy supplied to the field coil will be converted to kinetic rotational energy of the package. As the field coil rotates towards the equilibrium position, the accelerometer reading acceleration along the line of the orbital radius will sense zero acceleration as the angular relationship between the field coil and the earth’s North/South polarity reaches 90 degrees relative to the maximum repulsion or attraction position. The coil polarity control circuit is designed to reverse the polarity of the field coil at this position, thus maintaining a repulsive relationship as the rotational inertia carries the coil package through the 90 degree position.
The timing of the polarity reversing switch is critical for maintaining repulsion; avoid dampening the oscillation, or allowing the package to continue increasing its spin velocity.
As power is applied to the circuit, energy begins to be converted thru linear and rotational acceleration to the gravitational and rotational energy of the package. Without empirical proof, I suggest that the applied coil circuit energy will be the sum of the energy translated to gravitational potential and rotational kinetic energy. That the linear force acting along the line of orbital radius will vary as the cosine of the relative field angle while the force translated to torque about the center axis of the coil will vary as a sine function of the relative field angle. The linear force will approximate the force at 0 degrees (maximum repulsion) times the cosine of the relative field angle. The force imparting torque about the center axis of the field coil will approximate the force imparting torque about the center axis of the field coil at 90 degrees (maximum torque) times the sine of the relative field angle.
Where the moment of inertia of the experimental package is known, Earth’s Electro-Magnetic Field Strength can be derived from acceleration and circuit power.
Attitude Determination and Control Applications On Board Earth Based Satellites:
Using the Earth’s magnetic field as a reaction field in attitude determination and control of earth orbiting satellites was first proposed in the early 1960’s. Current applications include the sensing of relative field angle to determine satellite attitude and the use of electro-magnets to maintain and change satellite attitude. The author believes that widespread application may be limited by the orbital perturbation that would result from earth field/satellite field interaction. Electro-magnetic attitude control, without an Electro-magnetic orbital maintenance regime, would require expenditure of thruster fuel to maintain orbital station. To make electro-magnetic attitude determination and control a viable application, a means of offsetting the orbital perturbation using electro-magnetic propulsive technology rather than chemical thrusters must be developed. Also, the mass and volume fractions of electro-magnetic attitude determination and control technologies must be brought to values where the advantages of the technology offset the mass and volume fractions required. A primary advantage of Electro-magnetic propulsive methods for this application is that it can be accomplished without mechanical components as required in momentum and reaction wheel technology or the fuel and valving required for thrusters. This resulting increase in reliability will serve as further incentive to apply electro-magnetic technology to attitude control.
Increasing Hyperbolic Excess Speed in departure from Earth’s sphere of gravitational influence:
By definition, for an Earth orbital escape mission, the Hyperbolic Excess Speed is the residual speed that remains as the space craft climbs out of the Earth’s gravity well. Simply stated, it approximates the rocket burn out velocity minus the escape velocity at the range of burnout.
It may be possible to increase the Hyperbolic Excess Speed by using magnetic repulsive force generated by propulsion coils aboard the spacecraft acting against Earth’s electromagnetic field. The repulsive force would offset the deceleration of gravity as the space craft moves out of the Earth’s gravity well. This offsetting force would leave more residual or “Hyperbolic Excess Speed” as the space craft leaves the gravitational sphere of influence. If the magnetic repulsive force exceeds the gravitational force, then this force would continue to accelerate the space craft, imparting additional mechanical energy. The additional specific mechanical energy conserved or imparted would approximate the applied circuit energy calculated as applied power times time.
By thoughtful design, repulsion can be maintained without using an oscillating polarity strategy (as described in the Space Based Experiment), thus maintaining constant space craft attitude.
Increasing Hyperbolic Excess Speed in Gravity Assist Maneuvers:
A number of deep space missions have used Gravity Assist to either increase or decrease the mechanical energy of the space craft. Although such maneuvers use the gravitational acceleration of the assisting planet to increase or decrease the specific mechanical energy of the space craft, the Hyperbolic Excess Speed of the space craft relative to the assisting planet remains unchanged. The reason for this is the relative velocity between the space craft and the assisting planet gained by the acceleration of gravity on the approach trajectory is lost to that same gravitational force on the departure trajectory.
By using Electro-magnetic propulsion, additional hyperbolic excess speed can be imparted when the assisting planet has a significant magnetic field. In this application the propulsion coil(s) are used in attraction polarity on the approach to the assisting planet. This increases the acceleration above that imparted by gravity alone. As the space craft begins its departure trajectory from the assisting planet the relative polarity is reversed and maintained in repulsion; offsetting the deceleration imparted by gravity, thus conserving the Hyperbolic Excess Speed of the spacecraft relative to the assisting planet by adding the energy imparted by the electro-magnetic system to the specific mechanical energy of the space craft.
Orbital Station Maintenance and Altering Eccentricity:
By using properly timed Electro-magnetic Impulse, in repulsion and in attraction, possibly combined with Electro-dynamic braking; total orbital energy and eccentricity of orbit may be altered. Conceptualization of this regime involves both magnetic impulse and dynamic-braking at specific points in the orbit including using the induced dynamic-braking energy to produce vectored magnetic impulse.
Total Orbital Energy, often referred to as Specific Mechanical Energy, has two components. The kinetic energy per unit mass and the gravitational potential per unit mass. The sum of these two variables equals the specific mechanical energy. In an orbiting object, when not acted on by any other force other than the gravity of the prime focus object, this Specific Mechanical Energy remains constant. In elliptical orbits this energy translates between kinetic energy and potential energy as described by Kepler's second law.
Introductory texts on Astro-dynamics teach that in most cases, a change in the Specific Mechanical Energy of a satellite is accomplished by imparting impulse along the velocity vector. This change in velocity translates to a change in the radius of orbit. By imparting impulse along the velocity vector the Specific Mechanical Energy can be either increased or decreased with the timing of the impulse relative to periapsis or apoapsis determining orbital eccentricity. Changes in apoapsis are made by imparting impulse at periapsis while impulse to change periapsis is imparted at apoapsis.
In both cases, the impulse either increases or decreases the total orbital energy by imparting a change in orbital velocity. This change in velocity is then translated to a change in gravitational potential by altering the semi-major axis.
It is proposed that changes in the orbital energy of the space craft can be made using Electro-magnetic technology; increasing the orbital energy by increasing the semi-major axis directly through repulsive interaction with earth field or decreasing orbital energy by electro-dynamic braking.
In a low inclination orbit, generating a magnetic field in repulsion with earth field will begin to impart impulse along the orbital radius, increasing the semi-major axis, thus directly increasing the gravitational potential component of the total orbital energy (Specific Mechanical Energy). By imparting magnetic impulse during the entire orbital period, or applying bit impulse relative to apoapsis and periapsis, the orbit can be stepped up and eccentricity controlled. If using conventional chemical propellants, stepping up the orbit is accomplished by increasing the velocity component, translating to gravitational potential, with the impulses timed relative to apoapasis and periapsis to control eccentricity. Experimentation with generating magnetic field in attraction to earth field may yield some surprising results. How will the circuit energy be conserved?
Electro-dynamic braking of the space craft will impart a braking force along the vector of orbital velocity, decreasing the orbital energy, and translated a reduction in gravitational potential. By timing the Electro-dynamic braking inputs relative to periapsis and apoapsis the orbit can be stepped down and eccentricity controlled. Using chemical propellants stepping down the orbit is accomplished by impulse opposite the velocity vector, slowing the spacecraft. The timing of braking impulse relative to periapsis and apoapsis will allow control orbital eccentricity.
All electro-magnetic induction processes are composed of three primary components; the excitation field, the inductor, and rate of change. The rate of change can be supplied by velocity of the inductor relative to the excitation field, the oscillation of the excitation field in the presence of the inductor, or combination of the two.
For those of you who may have had the opportunity to empirically experience the fundamental physics of induction through experimentation with a simple hand crank generator, that lesson showed the relationship of circuit load to cranking force, and can be extrapolated to the inductive braking of a satellite or asteroid.
In the hand cranked generator, the permanent magnet supplied the excitation field for the induction process. The rotor, turning a coil through that excitation field, supplied the “Rate of Change” required for induction. The faster the rotor was turned the higher the voltage that was developed across the leads of the generator. When there was no “load” across those leads the generator was very easy to crank even though the “potential” or “voltage” was still being developed across the leads. But when a load, such as a light bulb, was placed in a circuit across the leads of the generator, that bulb created a “load” in the circuit. That “load” was the energy being dissipated in the bulb through resistive heating and light production. Supplying the circuit load with energy required greater force in cranking the generator. The force applied to the handle of the generator was converted to torque by the lever arm (handle) which spun the inductors (coils) at near right angles relative to the magnetic field (excitation field) of the stator. The energy required to spin the rotor was nearly equal to the energy being dissipated by the circuit “load”. This is a good empirical example of the law of conservation of energy.
In Electro-dynamic braking applications the solar or planet field will provide the excitation field while coils aboard the spacecraft, or the spacecraft/asteroid body itself, serves as the inductor. The spacecraft/asteroid velocity cutting the flux lines of the Solar or Planet field will supply the rate of change component. This induction process will generate a braking force as is inherent in any electro-magnetic induction process. The induced energy will then be dissipated through circuits designed to generate heat for radiative dissipation, conversion to vectored propulsive impulse, or stored for peak power/subsystem applications. Applications of Electro-dynamic braking will include adjustments in semi-major axis and eccentricity; as well as braking to orbit in planetary missions.
Because the orbital velocity of a satellite or asteroid is so high, significant voltage can be developed even though the excitation field may be very weak. The “tether” experiments flown on the space shuttle clearly indicate the validity and potential application of this technology.
Imparting Orbital Escape Energy:
Escape speed, as given in reference material, gives the escape speed at the surface of the body referenced. This escape speed decreases as the radius of orbit increases. If an orbiting spacecraft is given continual magnetic impulse to step up the orbital radius, at some point, the orbital velocity of the spacecraft will approach and then exceed the escape speed at range, thus allowing the spacecraft to “escape” the gravitational sphere of influence of the prime focus body (planet or Sun). Using such a method, a satellite may be given Excess Hyperbolic Speed, not by imparting additional velocity, but by imparting additional gravitational potential until the Specific Mechanical Energy of the spacecraft exceeds that needed to escape the gravitational sphere of influence of the prime focus body.
By initiating Electromagnetic impulse at high power in low earth orbit, very high Excess Hyperbolic Speeds may be obtained.
Braking to Orbit:
A space craft approaching Jupiter, or other target body, may have too much energy to establish orbit. By using Electromagnetic Propulsive methods it may be possible to alter both the magnitude and vector of the approach velocity; thus giving an alternative to using chemical propellants or atmospheric braking as the sole methods of reducing the energy of the spacecraft so that it can be captured by the intended prime focus body.
Deep Space Propulsion and Navigation:
In propulsion and Navigation, Electro-magnetic Propulsion will use electrical energy generated by an on-board power source to generate electromagnetic field(s) which will impart impulse through combined interaction with the magnetic fields of the Earth, Sun, Planets, and eventually, galactic fields.
Force vectoring will be obtained by very precise control of field strength, field angle, and action time relative to those reaction fields. Vector control, when within the magnetic sphere of influence of multiple field sources; will utilize the relationship of reaction field range, reaction field angle, and time span of power input to sum the force vectors from two or more reaction fields to obtain the desired net force vector. An example would be to act in repulsion of earth field for a fixed time at a fixed power and then act in attraction to Sun field for a fixed time at a fixed power. The net vector force would be the vector sum of the two forces. Because of the cosine relationship of repulsion/attraction force to the relative angle between the propulsion coil(s) and fields of the Sun, Earth, Jupiter, or other reaction fields, and if those fields are offset at a substantial angle relative to each other; effective impulse vectoring can be accomplished.
The Delta V imparted by a chemical rocket is limited by the attainable exhaust velocity of the rocket and total fuel mass available. In an Electro-magnetic Propulsion System capable of generating extreme field strength, the Delta V imparted will be limited only by the amount of available applied electrical power and the time span that power is made available.
In all cases, the minute field strengths of the reaction fields at range become usable when the propulsion system is capable of generating extremely strong fields or, in the case of dynamic braking, the induction circuit is capable of maintaining extremely high acceptance when dissipating high power.
The use of Electro-magnetic Propulsion may negate the requirement of waiting for opportune alignment of Jupiter and Saturn for use in Gravity assist maneuvers. The launch windows now continuously open by the ability to use Solar Field in repulsion to give the space craft the kick up that would otherwise require a gravity assist trajectory or maneuver.
A most important ramification of Electro-magnetic Propulsion and Electro-dynamic Braking in Space Applications may be the fundamental change in the logistics of asteroid deflection. This technology will negate the need to carry chemical fuel mass to the asteroid for the purpose of supplying impulse. It will allow mankind to use the orbital energy of the asteroid itself as the prime source of energy for deflection through an integration of Electro-dynamic braking, vectored electro-magnetic impulse, and Newtonian Propulsion Systems that use scavenged mass from the asteroid and accelerate it using propulsion coils. Perhaps, it will cause a re-evaluation of the decision to use nuclear explosive deflection and fractionation as the preferred approach to this impending challenge.
By Mark J. Carter