Technology Primer (20p) — H.Mk.0: Reusable Interorbital Transport. 191120.v06
M. J. Wegener. https://h-industries.io
Public GDrive — Original Post: 191120.v06 / Medium Rehost: 200105.v06
Introduction — Concept Development — Interorbital Transport Service — Launch Mechanism — Satellite Construction — Swarm Construction
H.Mk.0 Electromagnetic Launch System — Launch Function — Launch System Algorithm — Concept Proof Chain
Replace interplanetary rockets with a low cost, high efficiency solution for routine freight transport.
To climb Mount Everest, a number of stages must be crossed, all with differing requirements. In order to reach base camp, a good jacket and hiking boots will see most travelers safely to their destination. However, after base camp, the requirements change drastically. The following stages of the ascent will require mountaineering equipment, weather resistant clothing and possibly an oxygen supply. Though these requirements are physically harsh and require significant preparation compared to the approach stage, they are consistent. Mountaineers may carefully plan and execute a successful summit, though the less prepared amongst us are still able to glimpse the rarified peak by outsourcing the issues. This does not reduce the risks or arguably the thrill of the adventure, it opens access to the world’s tallest peak to a greater number of people by removing specialised effort. Just as a sherpa is able to bring your equipment to base camp, a seasoned mountaineer will guide you on the journey over the summit. For each stage of the journey, the outsourced service is tailored to the requirements of the environment. The same logic must now be applied to crossing the cold void of space.
This paper is a technology primer for design and operation of a reusable interorbital transport system. This system will move any material in freight containers between satellite swarms in geosynchronous orbits around any planet or celestial body. The idea was developed over four years; concept design was built up to solve issues of functional integrity before the system was validated through academic research. The system was designed using existing components and consumer grade materials before looking at research grade componentry. It is recommended to read through the content briskly at first as each new element is stated then explained in the following sections to support the next dependent element. Review this paper with a second reading for clarity or visit the website and YouTube for more.
This paper explains concept development and operational context before detailing the launch system. The capability of the Mk.0 satellite design constructed from current consumer grade materials is used to define the operation of an interorbital transport system. This is the public release section of an evolving research and design paper containing satellite design, swarm design, operational procedures and failure mode analysis. This paper is the first publication of any vector additive electromagnetic propulsion system and presents a new use case for satellites, they are just the delivery mechanism for a number of other existing technologies.
Innovation doesn’t necessarily require doing something new, it can be as simple as using what we already have in a new way.
Space is primarily a cold, empty void. The vacuum reaches 2.7 Kelvin at its coldest while soaring to thousands of degrees near celestial objects, this represents an opportunity as much as a constraint. The effects of the intense radiation environment, drifting particulate matter and large celestial objects are all issues that must be contended with if we are to leave our planetary confines. These elements combine to present significant challenges in many different circumstances however they are at least consistent, much like the conditions faced crossing the summit of Everest.
The void of space is a vast empty expanse, free of the fiction causing environmental elements that generate tremendous heat on atmospheric re-entry and significant reductions in the velocity of moving objects. As the void lacks abrasive, drag inducing molecules, there is every chance that an object will continue on its set trajectory indefinitely until it runs into something. This presents a significant energy advantage to any form of travel.
Superconductive materials are limited in terrestrial applications due to their temperature dependence. Significant research is invested in developing high temperature superconductors but in the vacuum of space this terrestrial requirement is not as relevant. Temperatures in orbit remove the refrigeration energy cost of superconductive materials and their coolant baths. This is a challenge for supporting components that must be maintained near room temperature but the issues of one system neatly presents solutions for the others. The development of electromagnetic propulsion systems for terrestrial use has been limited by these material properties and environmental conditions. To beat air drag, an energy cost barrier must be overcome. Normal, resistive conducting, components are not able to achieve the required energy densities or output at a size that allows for a viable design. By utilising the lack of air drag and the freezing temperatures of the void, significant progress is being made on electromagnetic propulsion.
The H.Mk.0 swarm satellite reusable interorbital transport system (RITS) enables a use case and application that is not in competition with any other electromagnetic propulsion system. The Propulsion Core is the sealed superconductive electromagnet at the heart of each satellite unit. It is a miniaturisation of electromagnet designs from CERN’s linear accelerators and current research. High density capacitor banks pulse power through a superconductive solenoid to create strong electromagnetic fields with maximum efficiency. All other designs are currently focused on making incredibly fuel efficient engines or in continuous low thrust systems that will allow us to send probes in to the depths of our galaxy. The H.Mk.0. system relies on a different application of many of the same principles. The energy will be generated, conserved and then consumed in explosive fashion using solar power to charge batteries which feed capacitor banks that release an explosive burst through a superconductive solenoid. This is more akin to loading and firing a gun. Each satellite in the swarm is triggered in concert to generate an electromagnetic propulsion wave that accelerates the launch plate & customer payload. The same procedure will be applied to an incoming payload to decelerate and place the customer in their desired orbital position.
Large rocket systems are effective at generating high thrust to escape the ground and reach orbit. Without higher fuel efficiency or frequent refueling, the high thrust capacity and time cost of a rocket system is not suitable for non-urgent journeys along stable shipping corridors. The Mk.0 Launch System requires 54 satellites in four layers for an operational swarm with one located at both origin and destination. Increasing the number of layers and density of satellites in each layer will correspondingly increase the acceleration and deceleration capacity of the swarm. The acceleration imparted will be limited by the deceleration capacity of the destination swarm.
Approximation is a first pass method for proof of concept, by demonstrating that one objects scale is an order of magnitude different, elements can be intuitively accepted or rejected as functional without a distinct derivation of proof. i.e. the problems are too small to have significant impact and you have a whole lot more power at your disposal. This solution was developed from consideration of the magnitude differences in particle acceleration and orbital mechanics. Just as large electromagnets can accelerate and control the trajectory of a powerfully charged particle, could they work on a larger object on a shorter journey, say a cargo container shipped along a corridor to Mars?
If accelerating a subatomic particle around a 27km track is to be accepted as a starting point for a cargo vehicle on an interplanetary arc then some proof by approximation is needed before digging into more complicated math. Concept development and the resultant proof will drive out the requirements to construct the components, which can be determined as either possible or impossible to manufacture.
Taking the principles of the idea, trimming the irrelevant parts and including those that could apply is simplistic but effective for building intuitive understanding. For the approximation of particle acceleration technology applied at an interplanetary scale, magnitude estimates are required to prove that most of the same rules will apply. Electromagnetic field quality is a key factor in maintaining accurate orbital trajectory for particles in an accelerator, thus the same will apply for passenger vehicles on interplanetary arcs. The Large Hadron Collider at CERN analyses particles at the quark level with a radius of just 4.3*10–18m orbiting in a loop of 27km (Russenschuck, 2010, Ch.1). When compared against cargo vehicles approximately 10 meters in height, traveling a mean distance of 78.39m km to Mars there is an appreciable order of magnitude difference. In our instance, the cargo is 5*10^0 m high on a journey of 7.839*10^10 m, a magnitude difference of 10^10. CERN’s 8.6*10^–18 m wide particle is on a 2.7*10^4 m trip around the LHC, a 10^22 magnitude difference. This leads to the conclusion that a number of factors can be acceptably lower while maintaining accurate interorbital targeting and trajectory.
The core requirements of the electromagnet assembly are energy input rate and superconductivity, needed to generate fields of the required strength. Superconductivity is a condition of virtually zero electrical resistance in a material as electrons unlock a shortcut beneath a specific temperature. Cooling of the electromagnet assembly beneath the critical temperature is energy intensive at ground level but less so in High Earth Orbit at the edge of space. This leads to another order of magnitude consideration, if the magnet strengths required are identical then it is impossible to embed an electromagnet assembly of the CERN quadrupole scale in one satellite. How about distributing it between many?
Acknowledging the decreased level of control, accuracy or field quality required does not reduce the energy input needed to potentially accelerate the object in question. This leads us to the conclusion that at a first pass, we need to build distributed electromagnet assemblies following the design principles of the LHC while optimising maximum potential strength and power storage capacity. The decreased field quality requirement is used later to determine how accurate the vectors created by the individual electromagnet assemblies in each satellite must be. Vector addition is adding small impulses together that combine into a unified force and direction. The same idea applies to the generation of propulsion vectors from a whole swarm of satellites with high field pulsed superconductive solenoid assemblies. If the components have already been built, only a conceptual model and correct assembly is required.
Interorbital Transport Service
This system will be used to transport freight along shipping corridors between planetary orbits. This service is slower than rocket systems but far cheaper operationally and across the system lifecycle. Customers are offered return transport services between swarms or a one way launch towards any desired location. Our first crewed missions to Mars and other locations should be supported by years of scientific monitoring, preparation and supply delivery.
Launch service providers will deliver customer payloads in crates on a docking vector to be collected on steel plates controlled by swarm units. This allows any payload to be delivered safely to orbit around Mars first, then any planet or orbital body. Last mile ground delivery services are not included.
By offering interorbital transport as a cheap and accessible service, satellite swarms will become the railroad of the modern era. This will give the opportunity of interorbital expansion to a wide range of customers. Increasing utilisation is the primary driver to reduce cost and enable other projects to begin.
Induction is an effect created by an electric current energising a coil of wire that generates a magnetic field and creates a reactionary current and resultant field in other objects affected by the applied field. Solenoids are a type of electromagnet, with the wire coiled around a conductive rod that focuses the electromagnetic field. Electromagnetic propulsion is the momentum imparted by the ‘pressure’ of the electromagnetic field pushing against the induced current in the object.
Magnetic fields can be generalised as expanding in a mostly spherical manner from the top of the solenoid. This means the electromagnetic push from each satellite will affect the satellites in a hemisphere above it to drive them forwards. The push from successive layers of satellites combines the imparted momentum in the same manner as coils in a spring. The plate is launched from the top layer of the swarm and the reverse of this process is applied by the receiving swarm.
Figure. Electromagnetic Launch Pulse Interaction
Each satellite is a tall slim cylinder clad in white ceramic tiling with a recessed black glass iris on top. The frame is constructed from steel, with the body built using four vertical beams and a number of horizontal platters to divide and protect component sections. Four hinged wings extend from the base, containing solar panels with copper veining in a leaf-like pattern. The underside of the wings will have copper heat diffusion veins that act as ablative material to protect the satellite when the wings are closed. The wings will be open during launch to capture the electromagnetic pulsewave using the steel frame. The wing locking mechanism will ensure that momentum is transferred to the satellite body with minimal loss or structural stress. Collapsible wings allows for efficient transport of multiple units together and for situational closure to prevent damage to the satellite.
The Propulsion Core and power storage banks fill the majority of the internal structure. The remaining space will be used for refrigeration, electronics and wiring. Multiple positioning propulsion cores will be used for intraswarm movement while a secondary consumable fuel thruster on the units base is used to move the satellite before swarm establishment.
The test swarm will be a single triangular pyramid of four units to refine the system control procedures before more units are added and four satellite units launched to Martian orbit, repeating the process. The swarm core will be established before addition of units in layer then below for layer expansion. Initial configuration will use a square pyramid structure with 2x2 units in Layer n, 3x3 units in Layer n-1 etc. This requires 54 satellites for a four layer operational swarm. The launch of multiple units to an orbital destination is all that is required for swarm establishment.
The intelligent organisational capability of each swarm will be goal driven to abstract low level controls as firmware, distributing minor adjustment decision making among satellites rather than requiring operations control center resources. The new swarm will communicate to the operations center as it travels, organises and signals readiness for cargo receival.
Figure. Mk.0 Swarm Architecture
H.Mk.0 Electromagnetic Launch System
Each propulsion core is capable of generating a high strength pulse magnetic field with the capacitor system supplying energy to deliver the electromotive force. At a minimised spacing between satellite layers and units, this allows the maximum transfer of force to each of the satellite above. Using Newton’s Second Law, F = MA, with both vertical force and cargo mass known, the maximum potential acceleration of the system can be found and thus the trip time to Mars in the optimal case. On a 2000 kg payload resting on the cargo plate this gives controllable acceleration and vector direction from the combination of launch pulses. In the loss-less vacuum of space, this should make the mean trip to Mars of 78.39 million km in 2 years. The trajectory will be graded based on cargo type acceleration thresholds, with organic cargo limited to 8 G while mechanical cargo is limited to 20 G or 50 G for bulk cargo such as construction materials and nonsensitive supplies. The reduction in utilised acceleration allows for correspondingly larger and heavier payloads to be transported, increasing further with swarm growth.
The system accelerates and decelerates cargo plates using high strength pulsed electromagnetic fields. The launch mechanism is an interaction of Faraday’s Law of Induction and Lenz’s Law resulting in the ‘skin effect’ that conveys the force through each layer of the satellite swarm to the cargo plate. This is the induction of a current within the surface of a conductive object inside a magnetic field, creating an electromagnetic field at the surface that opposes the direction of the inducing field and thus pushes against it, creating motion. Lowering the material resistances increases induction and thus momentum.
Commercial maglev trains have operated since 1984 using the laws above and the system detailed here uses the same principles. One method of operating maglev trains is for the rail to have a single consistent current and polarity while the train itself provides power to the opposing electromagnets as required for lift, acceleration and deceleration. Just as the rail is pushing the train away (and vice versa) in its active state, held neutral by gravity, the same applies to satellites cooperatively pushing off each other. The satellites alternate between magnetic attraction and repulsion to maintain a constant spacing for grid arrangements. With stacked layers of satellites each pushing upwards in sequence, the layers act as individual coils of a spring to create the propulsion pulse. In a square pyramid, each satellite in the layer above is supported by four satellites beneath, with the combination of electromotive force vectors from each lower satellite directing the facing and trajectory of the satellite above. The vector combination upwards through each layer’s maglev maintained spacing allows the propulsion pulse wavefront to be adjusted by individual satellites for directional control of the resultant vector. This vector addition allows for accurate destination targeting and consistent launch speeds while accommodating variances in cargo weight distribution or error correction for misaligned satellites.
As each layer exerts force on the layer above, the reaction pushes against the satellites own imparted velocity from the layer below. With the concurrent triggering of external coils for a unified field across the swarm to break the orbital inertia, the foundational layer of satellites in the launch swarm must trigger their thruster assemblies to resist the magnetic repulsion and maintain orbital location.
As the swarm is moved by the reaction of the individual against the anchored magnetic mass of the whole, the foundational layer must respond to the reaction force exerted by the layer above. This anchoring acts to drive each of the subsequent layers upwards from their starting position. The next layer is then pushed upwards by the field generated in the foundational layer and induced field generated by the skin effect in the structure of the satellite wings and base. As this effect is triggered at the same time through each of the swarm’s layers, the layers expand in preparation for the inner coil’s acceleration impulse. The triggering of the acceleration impulse in the launch sequence sees each layer effectively pushed by the sum of the layer’s below initial momentum, before the top most layer with the cargo transport unit is ejected along the intended flight vector.
Figure. Launch Pulse Force Reaction
Due to the total lack of air resistance, the trajectory suffers no frictional velocity losses on its journey. The flight path failure points are disruption during the launch pulse generation or gravitational pull affecting the final trajectory. In particle acceleration applications, magnetic beam rigidity is a measure of the force required to deflect the beam from its directed path. A similar factor is considered here to ensure that no magnetic interference will impact the launch vector. As the initial swarms are in an approximately geosynchronous orbit around Earth and Mars, the system will primarily conduct launches from each planets night side to eliminate solar heating and utilise the native vacuum temperature to reduce the refrigeration requirements, allowing energy to be conserved for the launch and post-launch procedures. The swarm harvests solar energy for launch pulses during the day and conveys power to swarm units without appropriate exposure. This will then be consumed during the night side launch window to maximise system efficiency.
The launch trajectories may bypass other orbital bodies to deliver the cargo to its intended location, a standard orbital transport procedure. Due to the difference in orbital periods between Earth’s yearly and Mars’ biannual cycle, there are a number of regular bypass manoeuvers that will be used. This allows for shipping lanes to be established and regulated, reducing risks as launch frequency increases.
Launch System Algorithm
The following sequence builds dependent assumptions to model a swarm of satellites moving cargo. It is structured by combining existing formulas in as no current research describes a physically defined time variable electromotive multivector function. The problem context is defined then substitutions are made for applicable factors before assessing against a validation criteria. In this case, the system is viable if the targeting accuracy of the launch system is better than the required accuracy of the cargo plates transfer vector to arrive at the appropriate destination. This is determined by the flight path vector error bound being less than the width of the receiving swarm launch pad. Speed, trip time & cost efficiency are secondary if the cargo misses its destination.
First, a static function will be defined then time dependent elements will be added to model power consumption, pulsewave generation and force application in a dynamic scenario. The resulting function can be modelled in spreadsheets for initial verification before numerical analysis software is required to give an accurate determination of error bounds and final system design acceptance.
Coulomb’s Law is used to determine the force on a charged particle within a field. It determines the force based on Coulombs constant k(e), the electromagnetic charges Q(f) & Q(p) and square distance (d).
F = k(e)*((Q(f)*Q(p))/d²)
Equation. Coulombs Law
Consider the charged particle Q(p) as an uncharged object in an electromagnetic field of charge Q(f) generated by a pulsed electromagnet. The pulsewave induces a charge in the uncharged object, which in turn pushes against the applied field. Charge (𝑄) is defined by the voltage (V) generated by the applied energy (J). One coulomb of charge is defined as one Amp of current (I) per second (t).
Q = V/J = I*t
Equation. Coulomb Charge Conversions
Define the dimensions of the satellite, then the pulsed electromagnet within and the uncharged object as the base of the next layers satellite or the cargo holding plate. Design the electromagnetic coil, capacitor system, refrigeration system, and solar charging capability. To check for functional realism, review supplier components data sheets and see what already exists on the market, this minimises manufacturing risk and raises project viability if research grade materials are avoided.
Next apply the environmental effects of 2.7 K, the orbital vacuum temperature, creates superconductive conditions in for the electromagnet’s material resistances. As superconductors are non-ohmic materials, the standard V = I R substitution does not apply. The maximum current of a superconductor is limited by the strength of the magnetic field. The strength of the magnetic field 𝐵(x) is in turn limited by the operating temperature 𝑇(x), the material factors of the superconductivity critical temperature 𝑇(c) and the peak field strength 𝐵(0) at 0 K.
B(x) = B(0)*(1- (T(x)/T(c))²)
Equation. Operating Critical Field Strength
For critical current I(c), the higher the field charge and temperature are, the lower the current will be. Following the Critical Field Strength equation above, critical current density is approximated as:
I(x)=I(0)*(1- (B(x)/B(c))²)*(1- (T(x)/T(c))²)
Equation. Critical Current Approximation
Release the stored power as a pulse close to the critical current through the superconductive electromagnet to create a charged electromagnetic field Q(f). Design will define the limit for charge generation to either the capacity of the electromagnet or the power storage system.
The field strength at any point above the electromagnet is determined by the point’s distance (z), coil radius (r), length (l) and the current (I) carried by the material. The Biot-Savart Law shows that stronger currents give stronger magnetic fields:
B(0) = (𝜇 /4𝜋) * (𝑟𝐼𝑙 /𝑧)
Equation. Biot Savart Law — Electromagnet Field Strength
The total voltage drop and the inductance of the capacitor system defines the time component of the current consumption rate (A/s). The delta between the fully charged voltage and the depth of discharge limit voltage gives the pulse transformer flow rate requirements. To use this energy consuming electromagnet at the maximum energy input rate, a power storage system must be designed then both systems optimised together. The electromagnet consumes energy as it engages and rises to the stable field point, this is the flat top of the power pulsewave. Energy is measured Watts (W), Joules (J) per second (s) and is the work component of the electrical voltage (V) and amperage (I).
Equation. Work Energy Formula
Substituting all the components above into the initial equation shows the pulsewave profile gives force (F) applied over time ∆(t), giving impulse to the layer of satellites above. The energy required for a change in momentum of a satellite in the second layer is the product of the pulse waves force over time applied to the satellite mass (m), which can be used to find the objects final velocity v(f). By assessing the stages of the launch pulse as a time stepped function, the force imparted at the start of each timestep can be used to calculate change in momentum from the objects initial velocity at that time v(i).
F = m(v(f) −v(i))/∆𝑡
Equation. Pulsewave Force Applied
Acceleration is the derivative of velocity, determined by the force applied to a mass. The launch acceleration is given by the velocity formula, relating acceleration to time, along the path defined by the sum of applied vectors. Due to the lack of atmospheric friction, the imparted velocity has no loss or drag components. Solve for four layers of satellites and a cargo payload then determine the error bound to give target accuracy.
Once the force is imparted to each object, it will continue along the resulting vector indefinitely as there is no frictional resistance that could affect the vector in the vacuum of space. This is the key element that allows electromagnetic propulsion to succeed once the constraints of atmospheric friction and gravity are removed. With the goal trajectory defined from point to point, apply factors for the orbital context. This requires the object’s orbital height, speed, spin and facing. Assume ideal facing and rotational conditions for now, include the target location as above, position and orbital vector. These conditions give the trajectory vector path while swarm dimensions give the accuracy requirements.
(O𝑟𝑖𝑔𝑖𝑛 𝑆𝑤𝑎𝑟𝑚 Orbital Vector) + (𝐿𝑎𝑢𝑛𝑐h Trajectory) = (𝑇𝑎𝑟𝑔𝑒𝑡 𝑆𝑤𝑎𝑟𝑚 𝑂𝑟𝑏𝑖𝑡al 𝑉𝑒𝑐𝑡𝑜𝑟) + (𝑇𝑟𝑎𝑗𝑒𝑐𝑡𝑜𝑟𝑦 𝑂𝑟𝑏𝑖𝑡𝑎𝑙 𝑉𝑒𝑐𝑡𝑜𝑟 𝐸𝑟𝑟𝑜𝑟 𝐵𝑜𝑢𝑛𝑑)
Equation. Flight Path Calculation
If 𝐿𝑎𝑢𝑛𝑐h 𝑆𝑦𝑠𝑡𝑒𝑚 𝐴𝑐𝑐𝑢𝑟𝑎𝑐𝑦 ≥ 𝑅𝑒𝑞𝑢𝑖𝑟𝑒𝑑 𝐹𝑙𝑖𝑔h𝑡 𝑃𝑎𝑡h 𝐴𝑐𝑐𝑢𝑟𝑎𝑐𝑦 = 𝑇𝑟𝑢𝑒 → 𝐷𝑒𝑠𝑖𝑔𝑛 𝐼𝑠 𝑉𝑎𝑙𝑖𝑑𝑎𝑡𝑒𝑑. = 𝐹𝑎𝑙𝑠𝑒 → 𝐹𝑖𝑛𝑑 𝑇h𝑒 𝐿𝑖𝑚𝑖𝑡𝑖𝑛𝑔 𝑃𝑎𝑟𝑡. 𝐼𝑚𝑝𝑟𝑜𝑣𝑒. 𝑅𝑒𝑝𝑒𝑎𝑡.
Equation. Concept Validation
The numerous additional considerations of reality are covered in the H.Mk.0 Internal Design Paper.
Concept Proof Chain.
Quantifiable and dependent steps for modelling proof of concept.
Define satellite external dimensions.
Define container and launch plate dimensions.
Define swarm architecture dimensions.
Define electromagnetic coil to capacitor storage ratio.
Determine available satellite internal volumes.
Define coil wire and winding dimensions.
Determine strand utilisation, iterate as needed.
Determine coil inductance, mutual & wire self inductance.
Determine coil strength capability and force output.
Determine coil energy rise requirement.
Define capacitor system dimensions
Determine Watt/hour output from volume & chemistry.
Does capacitor Watt/hour capacity exceed coil energy rise requirement?
Determine pulsewave profile from ramp times & power consumption for flat top pulse length.
Determine field strength in force at top of solenoid, z(0).
Determine field strength & inductance in material at corner of satellite wing above.
Determine induced field in material then reactionary force against applied field for time.
Combine power / field strength / pulse wave / acceleration vector as electromotive model.
Iterate design to maximise pulse outcome.
Define cargo payload mass.
Apply electromotive function to each layer then cargo plate and cargo payload mass.
Define orbital location.
Determine orbital speed & rotation vector.
Apply planetary vector paths.
Determine transit vector paths.
Determine flight time.
We already have drone swarms and high strength pulsed magnets for linear accelerators. The design can be proven with 40-year-old superconductors and YouTube has dozens of videos of the ‘Ring Launcher experiment’, giving a physical demo of the launch mechanics. The orbital context removes air drag and the energy cost of cooling the pulse magnets giving a better system efficiency.
Space is more accessible than ever, cost efficient cargo transport is next after cost efficient orbital access.
Colonisation of Mars is inevitable, so interplanetary transport is a necessity. Save rockets for people and build out freight transport at maximum time cost efficiency. It is as inconceivable today as it is inevitable tomorrow and this system can be implemented in five years using a low cost and mass producible design. Manufacturing satellite units at scale will allow rapid expansion of an interorbital transport network with multiple swarms driving down the cost of transport on busy commercial routes. Stockpiling of supplies and regular delivery of foods, construct materials and other equipment is essential for maintaining a permanent presence off Earth.
Solving interorbital transport at the same time as making orbital entry accessible will revolutionise commercial space opportunities. Most businesses are focused on servicing our planet, why not expand services to several destinations?
Thanks for reading, @h.mjw
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#BonusPage — Advanced Use Case
After successful delivery of the initial business plan, the next step for H. Industries is to facilitate the capture of asteroids in partnership with exploration geology businesses by providing the asteroid receiving harness. A large swarm can slow the trajectory of ferric asteroids by applying the same cargo receival procedure to decelerate the asteroid then slowly halt its trajectory. The swarm can deposit the asteroid in an orbit of the customers choosing. By enhancing the capability of the terrestrial swarm and utilising the Propulsion Core in an alternative induction mode, the ore of ferric asteroids can be removed. This function is enabled by the difficulty of a material to dissipate heat in a vacuum, allowing ferric materials to be increasingly heated by induction. This simply requires that the energy input exceeds the materials heat loss. Technical advances in solar energy conversion will provide efficient sustained induction however the timeframe required for these advances can be mitigated by the application of more swarm units. Once an asteroid is stabilised in a terrestrial orbit, the swarm will reconfigure in a net to surround the asteroid by using the magnetic tethering between satellites. This swarm net will be used as a distributed induction furnace, using the Propulsion Core in each unit to heat the ferric ores within the asteroid. Heating of the ores will result in expansion, leading to cracking of the asteroid once sufficient expansion pressure builds.
Cracking of the asteroid and expansion of the ferric materials will allow the ores to be magnetically separated from waste materials. Separation of ferric materials will be followed by intensified induction applied to heat the ores for removal of oxides and impurities to isolate raw iron. This will allow raw iron to be extruded from the ferric material mass in a plastic fashion by applying directed magnetic impulses to fashion a contained liquid pipeline from the condensed ore material to a storage sphere. Raw iron will be used for its low melting temperature, ability to be easily cast using extrusion moulding techniques and the lack of oxidation in orbit. These factors remove the key weaknesses of raw iron as a base construction material in terrestrial applications, where it’s compromising oxidation and low strength to weight ratio make steel the preferable terrestrial choice.
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