JuanCastillo29/space-mission-intelligence-data huggingface dataset:
This study includes the use of infrastructure as part of the total mission. For two concepts whose performance is otherwise equal, the one requiring less infrastructure is preferred. Consider the classic interstellar concepts like Daedalus or Forward's microwave sails which assumed the preexistence of a substantial in-space infrastructure. Which of those divergent approaches would require less time and energy to build, assuming all other factors were equal? For example: will it be easier to mine He3 from the atmospheres of the gas giants to support fusion propulsion, create a dedicated antimatter factory to support antimatter rockets, or to build TW lasers and 1000 km diameter Fresnel lenses? And in the course of answering those questions, which requisite technologies might be revealed to be more crucial or broadly applicable?
Recall the distinction between the era of precursors and the era of infrastructure. In the strictest sense, the era of precursors use established ground-based infrastructure and propulsion methods that are mature enough to proceed to trade studies. Hence those dependencies are not measured here. For true interstellar missions, however—the ones reaching to the nearest starts—new infrastructure is required. Regarding the era of breakthroughs, those concepts will fall into either the precursor or infrastructure era, on a case-by-case basis, depending on the scale of mass and energy required. Those concepts might need to be advanced to TRL 3 before such a determination can be rendered.
There is a split in the era of infrastructure concepts; those concepts that involve power beaming from the surface of the Earth (like StarShot), and those concepts that assume the use of preexisting in-space infrastructure (like Daedalus and Starwisp). These require two different assessment methods.
The fundamental units for comparing infrastructure dependence are mass, energy, and time. In principle, larger and more massive spacecraft will require more time and energy to build. The same is true for the systems that beam energy to the spacecraft.
For the ground based infrastructure systems, only the energy to build the beam system is measured, not mass nor the time to build that beam. This is because Earth-based construction will likely be considerably faster than inspace construction. For the in-space systems, both energy and time are measured. For the in-space scenarios, two factors need to be assessed, 1) predicting what infrastructure will be available for use over the future years, and 2) then estimating how much of that infrastructure is used to support the development of the mission.
These analyses are more relative than deterministic. In other words, the assessment can distinguish which mission plans will require more or less infrastructure, but will not be able to accurately predict how much time and energy that process will require.
One further note: It is reasonable to consider that the same kind of laser array considered for propulsion might also be the primary source of power delivery for the entire infrastructure. The assessment methods to include that possibility have not yet been established.
For those concepts which aim for true interstellar distances (≥ 270,000 AU) and that assume their beaming infrastructure resides on the surface of the Earth, it is already possible to compare those systems against each other in terms of financial cost, instead of the more general measure of energy used throughout this study. To compare these concepts in energy terms, a conversion factor of $0.06/kW-hour, or 60 MJ/$ will be used. The economic theory behind this conversion [92] would be a lengthy discussion. Put simply, because energy is such a fundamental input to almost everything else made and used in civilization, the “real” cost of energy (in terms of how many goods and services a ‘unit of energy’ buys) changes very slowly with time. Small errors in the figure chosen as a conversion factor is not very serious in terms of comparing propulsion systems, so long as it is used consistently. The estimates for capital cost of a given beam can then be converted to energy terms using that conversion factor. If inventions change the capital cost of a given type of beam (lower $/Watt), then that is essentially equivalent to improving the “propulsion efficiency” of that type of beam, in that the lower $/Watt of beam power shows up in the comparison charts as a lower “total energy” (because of the reduced capital cost of the installation, expressed in energy terms).
It is conceivable that some Earth-based infrastructure might also have in-space components (hybrid concepts such as ground based optical beams with space-based focusing optics, or space-based particle beams with their
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Millis 2018 Grant NNX17AE81G_for_CR.docx power fed from the ground, are possible).

The cost to create these new elements can be estimated by the launch costs ($/kg, J/kg) required to get that hardware into its orbital position.
raw materials (including propellants) • Processing the raw materials into usable stock [100 MJ/kg aluminum?] • Manufacturing components from stock • Constructing objects from components and stock • Transporting items within the infrastructure Infrastructure Use Specific Energy, Esi (J/kg): For the analyses that follow, the specific energies for each of those functions will need to be estimated during stage-II. The sum of those specific energies will then represents the
Millis 2018 Grant NNX17AE81G_for_CR.docx

pg 43 of 69 total Infrastructure Use Specific Energy, Esi, to be used when calculating the use of infrastructure.

It is assumed that theses values might change over the decades of progress.
Infrastructure Available Power, Pia (W): The other key descriptor for the infrastructure will be Infrastructure Available Power, Pia, whose values will significantly change over the decades of progress, as provisionally indicated in the 3rd row of table 10. That factor will be used to determine how quickly the infrastructure energy can be expended, to estimate how long it takes to build things in space. Ideally, it would be more accurate to create a measure of infrastructure capacity, (kg/yr), from the estimates for the capacity for each of those functions. In the absence of that more detailed information, however, the more basic measure of Infrastructure Available Power, Pia (W), will have to be estimated as used for estimating how quickly things can be built with the infrastructure. Note that the only estimates obtained so far for capacity are for mining (kg/yr) for the 2030 and 2040 decades.
In principle, larger and more massive spacecraft will require more time and energy to build. The same is true for any beaming system for the mission. Thus, the first step of assessing infrastructure dependence requires knowing the total mass of the spacecraft and any supporting launch (beaming) hardware (whose calculations are described in Section 8.4):
• Spacecraft Launch Mass, Msl, (kg)
• Base Hardware Mass, Mb, (kg)

Spacecraft Hardware Build Energy, Esc (J): This is the measure of how much energy is required from the infrastructure to build the spacecraft. It is the product of the Spacecraft Launch Mass, Msl, (kg) and Infrastructure Use Specific Energy, Esi, (J/kg).
Base Hardware Build Energy, Eb (J): This is the measure of how much energy is required from the infrastructure to build the base hardware. It is the product of the Base Hardware Mass, Mb, (kg) and Infrastructure Use Specific Energy, Esi, (J/kg).
Total Infrastructure Energy Expense, Eti, (J): This is the sum of both the spacecraft and base hardware build energies, representing the total amount of energy needed to prepare the mission. The propulsive energy to perform the mission is discussed in Section 8.4.
The next question is: "how long will it take to build those objects?" The crude model used as a starting point is to divide the Build Energies by the Infrastructure's Available Power, Pia, (W), and convert the answer from seconds to years, whose sum is the variable named Mission Development Duration, Tdev (yrs).
Spacecraft Build (develop) Duration, Tscd (yr): This is the measure of how long it will take to build the spacecraft, calculated by dividing the Spacecraft Hardware Build Energy, Esc, by the Infrastructure's Available Power, Pia, (W), and convert the answer from seconds to years.
Base Hardware (develop) Duration, Tbd (yr): This is the measure of how long it will take to build the Base Hardware (like laser arrays), calculated by dividing the Base Hardware Build Energy, Esc, by the Infrastructure's Available Power, Pia, (W), and convert the answer from seconds to years.
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Millis 2018 Grant NNX17AE81G_for_CR.docx This is part of the more deterministic portion of the analyses.
trip time. This is not necessarily true for other systems.):
⎛
⎞
ΔV = ln M se +M pay +M p
M se +M pay
⎝ ⎜ ⎜
⎠ ⎟ ⎟ Eq. 8 And where:
Mse = Empty Mass of Spacecraft associate with Propulsion System Specific Power (kg) Mpay = Mass of Payload (kg) Mp = Mass of Propellant expended (kg)

From these, the Spacecraft Launch Mass, Msl (kg)—a value used in other portions of the analysis—is:
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M sl = M se + M pay + M p Eq. 9 Acceleration can be integrated for changing mass ratio, and for relativistic cases, should be. An approximation used here is based on average acceleration. Average acceleration, ā, (if not artificially limited to reduce loads on the payload), is then simply thrust, F, divided by the spacecraft empty mass, payload mass, and half the mass of the propellant as shown:
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a = F M se +M pay + 1
Eq. 10
2 M p For a flyby mission, the relationship between the total mission ∆V; acceleration (ave), ā; acceleration time, Ta; acceleration distance, Da; coasting time, Tc; mission distance, Dd; and trip time, Tt; follow these relations:
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Ta = ΔV
a Da = 1
2 a Ta
Eq. 11 Tc = Dd −Da
ΔV Tt =Ta +Tc Recall the other mission time parameter, Data Transmission Duration, Tx, discussed in Section 7.2.2, and 7.3. For orbiter missions, one must also account for braking time and braking energy, both of which can be significant. For that reason, orbiter missions often use a different propulsion method for braking. The extension from the above equations for type IP-OM systems is straightforward—half the delta-V available to accelerate, and half to brake. If another type of system is used, compute the braking time as appropriate for that system, and add it to the mission time.
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This sub-category of IP-OM types refers to systems like nuclear electric propulsion, where a separate power source (e.g., nuclear reactor) supplies energy to a rocket to expel an inert propellant (e.g., Xenon ion thruster). These are usually treated in terms of specific power (kW/kg) or alpha (kg/kW), and specific impulse (or effective exhaust velocity).
However, in these cases, “propellant specific energy” doesn’t fit. One cannot turn the energy into thrust without the reaction mass (propellant). A general assumption can be made that one runs out of propellant before running out of energy. Therefore, treating the propellant capacity as the limiting factor, there may indeed be a meaningful specific energy to define. Still, in any given case, the achievable acceleration and ∆V are usually clear from the particular apparatus in question.
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This type refers to systems which receive energy from external sources but carry their own reaction mass, such as solar-electric or solar-thermal rockets, or laser-illuminated electric or thermal rockets. The initial analysis provided here addresses the case of ground-based beam infrastructures which draw their power from the electrical grid. Their calculations for the propulsion energy also applies for the case of space-based infrastructure, but the additional cost to create that space-based hardware is explained in Section 8.3.3. Recall that the parameters that need to be calculated from the specifics of these types of propulsion include:
• Total Propulsion Energy, Etp (J) • Spacecraft Launch Mass, Msl (kg) • Energy Cost of Beam, Ebhw ($, converted to J using a J/$ scaling factor)
• Base Hardware Mass, Mb (kg) (For in-space infrastructure) [not yet addressed]
As these are still rockets, the equations of Section 8.4.1 still apply as far as computing trip time.  However,  now the jet velocity, Ve, is no longer determined by the energy intrinsic to the propellant as in type IP-OM, it is instead an adjustable parameter (within the limits defined by a given thruster technology; some thrusters have a narrow range of exhaust velocities, some a wide range). A given mission design therefore needs to specify the specific energy imparted to the reaction mass (alternatively, the exhaust velocity and thrust efficiency, which is equivalent, eq 5). In general, these systems are usually employed with total velocity change (∆V) which is comparable to or less than the exhaust velocity because the mass of power receptions and thrusters are often high enough to be incompatible with high mass ratios. Efforts to circumvent those limitations are ongoing.
The Propulsion System Specific Power Psp of these systems is usually dominated by the power supply, and affiliated systems, sometimes mistakenly neglected in naïve analyses, such as power conversion machinery and radiators for waste heat.
Furthermore, acceleration has to be analyzed in two phases: 1) when the beam or solar radiation fully illuminates the receiver at the limit of the power capacity of the spacecraft—in which case, it is the spacecraft which determines the attainable acceleration, and 2) when the spacecraft is too far to intercept full power of the beam, in which case the beam’s divergence causes available power to drop off with distance. A similar case exists for solarelectric propulsion, when the falloff of sunlight (1/r2) drops below the limit of the power capacity of the spacecraft.
The process specified here is for ground-based beam systems, where the beam's figures of merit are usually discussed as $/watt of beam power, and $/square meter of transmitter aperture. Thus the first generally relevant figure of merit for beams is:
Beam Source Cost Scaling = Cost of Beam Source $( ) Power of Beam Source W
Eq. 12
( ) Where "Beam Source" is the energy supplied to the beam, in terms of the beam's output power.
Since beams are, very roughly, characterized by a certain cost per unit area of aperture, and twice the aperture radius gives half the divergence angle, but four times the area, that implies a beam with half the divergence angle has the “aperture contribution” to beam cost roughly four times higher.
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Note that it is really the divergence angle of the beam that controls what the beam can do—different beams have different divergence properties that scale differently with aperture. Therefore, to define a figure of merit which is suitable for comparing both optical and particle beams, we want a figure of merit that is based on divergence angle. For optical beams, divergence angle is proportional to wavelength, and inversely proportional to the radius of the aperture. The beam director cost factor for an optical beam would be: Beam Director Cost ÷ Aperture Area, ($/m2).
But that doesn't work for other types of beams, including synthetic sparse apertures and particle beams, especially beams that include a degree of self-focusing. Instead, the second generally relevant figure of merit for beams is defined here as:
( ) × Beam Director Cost $( ) Eq. 13
Beam Director Cost Scaling = Divergence Angle radians
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Millis 2018 Grant NNX17AE81G_for_CR.docx Where "Beam Director" is the aperture-scaled hardware for focusing the beam.

The utility of that rather odd factor is clearer when put in terms of cost:
⎛
⎞
⎛
⎞
Beam Director Cost Scaling radians
⎜ ⎜ ⎜ ⎜ ⎜
⎟ ⎟ ⎟ ⎟ ⎟
⎝ ⎜
⎠ ⎟
$
Cost ($) =
Eq. 14
( )
Divergence Angle radians ⎝
⎠ It is obvious that a smaller divergence angle (tighter beam), gives a higher cost. So the divergence angle has to be in the denominator. Because of the scaling being such that, very approximately, half the divergence angle tends to be four times (or more) the cost, not twice. Given those two factors for a given beam technology, other beams of higher or lower power, or smaller or larger divergence angle, can be roughly estimated.
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This type refers to systems in which both the reaction mass and energy are supplied from external sources, such as solar sails, solar-wind magnetic sails, laser-pushed lightsails, and particle beam pushed magnetic sails. It also covers “pellet runway” concepts in which pellets, prepositioned ahead of the spacecraft, contain an energy source that the spacecraft uses to accelerate those pellets as a reaction mass (such as fusion-fuel pellets, e.g., Jordan Kare's "Bussard Buzz Bomb," an idea not published in journal form, but discussed in [98 p.112]).
While the details of each such system are quite different, and it takes some careful consideration of each system to fit it in to the overall comparison framework, they can be compared to each other, and to other systems, with the following approach.
As discussed in “Type RP-OM,” a given power transmitted by the beam, Pgb (W), and a given divergence angle, Θ  (radians), is characterized by a beam source cost scaling factor ($/W) and a beam director cost scaling factor (radians/√$ ). A given beam also has a given speed of the energy transmission, Vb (m/s), which is the speed of light for photons and less for particle beams. For particle systems, the speed of the beam can be varied during the mission which is a powerful tool for optimizing total energy used.
The beam in turn pushes on something, called here the “receiver”—which has a characteristic ‘Areal Density’ (that is, mass, per unit area of the receiver, kg/m2), and a ‘Flux Limit’ (the maximum power per unit area of the receiver, usually set by thermal limits of the materials, W/m2). When the receiver is saturated, operating at its power limit, we can define the now-familiar Propulsion System Specific Power Psp for this situation:
)
(
)

Receiver Areal Density kg/m2 As in more conventional forms of propulsion, the specific power in turn controls acceleration, while the beam is saturated. In the general case, however, there are usually two distinct phases of flight, 1) limited by the receiver, in which the receiver is saturated and taking all the power it can handle, and 2) limited by the transmitter, in which the receiver area (and mass) is determined not by power limitations, but by the divergence of the transmitting beam. For any given concept, it is usually clear how to handle these to develop overall trip times. However, the optimization of trip time for a given set of technology assumptions is not always straightforward. In order to be able to handle the
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variables, a concept should list both the areal density (kg/m2) and the flux limit (W/m2) of the receiver, because
some calculations need the parameters broken out.
Acceleration (Thrust/Mass) is then integrated over time to get velocity change. While in a realistic situation, the thrust tapers off as the beam diverges and less and less power is intercepted by the receiver, the equations have the same form as if the beam simply had a finite range. For example, Lubin has determined that for photon sails, the maximum velocity change is 1.4x the velocity obtained if the beam were cut off when the sail no longer intercepts the full beam (Beam Range Limit, Dbl) [99]. Therefore, while there is a correction factor (of order less than two), the velocity gain becomes:
⎛
⎞
ΔV = Beam Range Limit Correction Factor × 2 F M sl
⎝ ⎜
⎠ ⎟ Dbl Eq. 16
Millis 2018 Grant NNX17AE81G_for_CR.docx

pg 49 of 69
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Where:
Msl = Spacecraft Launch Mass (= Payload + Spacecraft Empty Mass,) (kg) Dbl = Distance of Beam Range Limit (m)

The Beam Range Limit, Dbl, is determined by the divergence angle of the beam and the aperture of the receiver.
For non-relativistic spacecraft speeds, the thrust for a photon beam is simply the power transmitted by the beam, Pbx (W), divided by lightspeed:
F = Pbx
c Eq. 17 For relativistic spacecraft speeds, one must reduce the value of beam power due to Doppler shift of the incoming beam.
The power to generate that photon beam, Pgb (W), is simply:
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Pgb = Pbx
Eq. 18
ηB Where:
Pgb = Power to generate the beam Pbx = Power transmitted by the beam ηB = Power efficiency of generating the beam
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−Vip ⎛ ⎝ ⎜ ⎞ ⎠ ⎟ Eq. 21
⎠ ⎟ 2Espη+Vip
⎝ ⎜
dt Where:
dm/dt = Participle Beam Mass Flow Rate into the spacecraft (kg/s) Esp = Pellet Specific Energy, where the pellets are both a reaction mass and carry energy (J/kg) η = Thrust conversion efficiency (%)

From this one can see that the mass flow required for a given thrust rises with the speed that the spacecraft encounters the pellets, until the power limit of the machinery that converts the pellets to thrust (a reaction chamber and nozzle) is reached. The beam cost in this approach is potentially much lower, because the pellet speeds can be much lower. However, the mission time is increased, because to the flight time must be added the time to pre-position the pellets; that factor limits the potential energy savings.
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This type refers to systems which receive reaction mass from external sources but carry their own energy. This is analogous to aircraft engines, where air is the reaction mass, and the energy source is the fuel. In terms of interstellar flight, this includes, for example, Bussard ramjets that gather protons and accelerate them with the help of onboard energy [101], and spacedrives that convert some form of stored energy into propulsive motion (kinetic energy) using as-yet-unconfirmed physics [82-87].
This category also includes “drag devices” such as magnetic sails (magsails) or plasma magnet sails. These devices are possible solutions to the challenge of braking at the destination. Without braking, the flyby time is very short. See table 8 in Section 7.2.1 for examples. After a flight time of decades, a flyby time of just hours seems disproportionate and would limit the fidelity of observations.
Therefore, the problem is not only one of how to get up to speed, but how to get rid of the speed. Drag devices—devices that serve the same purpose in interstellar flight as do aerobrakes and parachutes in planetary exploration—in principle dissipate the kinetic energy of the spacecraft against something else (usually, the ionized gases in the interstellar medium).
Finally, there is a class of potential “plasma wave” drive concepts in which traveling waves are launched in to the interstellar medium at a velocity far below the speed of light, and the resulting reaction force propels the spacecraft [102]. These are usually low specific power (low thrust) drives but they use the surrounding medium as reaction mass. Such an ability is similar to the goal of a spacedrive, but using existing physics.
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Depending on the nature of the reaction mass, the analysis methods vary. For things like plasma, interstellar protons, or stellar wind, analogies to aircraft propellers can be used. For new breakthrough propulsion physics (devices that can be viewed as using inertial frames or the properties of spacetime as an effective reaction mass), then other techniques are suitable.
Plasma Reaction Mass:
For concepts that interact with plasmas or interstellar protons, the fundamentals of conservation of energy and momentum result in the ‘propeller equations’ familiar for propeller and air-breathing jet operation within an atmosphere can be used:
⎛
⎞
P = 1
dm
2 −Vsm