JuanCastillo29/space-mission-intelligence-data huggingface dataset:
Fission Nuclear Fission Pulse Propulsion 1950 Teller-Ulam IP-OM Fission Fragment Rocket 1988 Chapline IP-OM Pulsed Fission-Fusion (PuFF) Propulsion 2017 Adams, R. Fusion Enzmann (3MT frozen deuterium ball to fusion rocket) 1964 Enzmann, 1973 Duncan IP-OM BIS Daedalus (Pulsed fusion, inertial confinement fusion) 1978 Bond 2 IP-OM Vista Inertial Confinement Fusion 1987 Orth IP-OM Project Longshot (Fission reactor pwr, fusion pulse propulsion) 1988 2 IP-OM Project Icarus (Pulsed fusion, inertial confinement fusion) 2011 2 IP-OM Continuous Electrode Inertial Electrostatic Confinement Fusions 2017 Sedwick IP-OM Fusion Driven Rocket (Direct Conversion) 2017 Slough IP-OM Gradient Field Imploding Linear Fusion Propulsion System 2017 LaPointe IP-OM Multi-stage fusion rocket IP-OM Antimatter ICAN-II, Positron catalyzed fission fusion 1998 IP-OM Antimatter-Catalyzed (pulse) Fusion (AIM star) IP-OM Antimatter - Matter Annihilation Propulsion Forward IP-OM
II.VI. Spacetime Warping

Gravitational Dipole 1963 Robert Forward 3 SP-XM Traversable Wormholes 1988 Thorne, Visser 3 SP-XM
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Millis 2018 Grant NNX17AE81G_for_CR.docx With the participation and co-sponsorship of Tau Zero via this grant, a workshop was held from October 3rd-6th, 2017, in Huntsville, Alabama, called the "Tennessee Valley Interstellar Workshop (TVIW)." The event brought in over 140 attendees, who spent three days hearing the most current and cutting-edge presentations by leaders in space development and interstellar flight and exploration fields.
Speakers were from the Breakthrough StarShot Initiative, General Dynamics, Johns Hopkins Applied Physics Lab, several NASA facilities, University of California, University of Washington, two United States Congressmen, and several representatives from the United States Air Force. Globally, speakers’ origins spanned from Sweden, to Germany, Japan, Italy, and Russia.
This is a continual and evolving forum to enable working discussions on interstellar research and exploration to encourage continued advances, and the publication of those advances, spanning education, technical research, societal facets, literature and cultures, and to enhance public attitudes and dialogue about interstellar exploration.

[Table: Table 5 lists the presentations and their authors, grouped by topic. Instead of showing the titles, short descriptors of the talks are given. The individual reports have not yet been published. Some will be published in a future issue of the Journal of the British Interplanetary Society. However, all the presentations can be viewed on line at: https://tviw.us/2017-presentation-video-archive/] [Table: Table 5.  2017 Workshop Lectures & Authors Mission Considerations and Overviews]

| Table 5. 2017 Workshop Lectures & Authors |  |
| --- | --- |
| Mission Considerations and Overviews |  |
| Summary of workshop Paul Gilster | [52] |
| How to assess interstellar challenges and prospects Marc Millis | [53] |
| Precursor mission studies with current technology Pontus Brandt | [32] |
| Exoplanet mission concepts from NASA JPL Stacy Weinstein-Weiss | [33] |
| Solar gravitational lens viewing of exoplanets - prospects Slava Turyshev | [37] |
| Solar gravitational lens viewing of exoplanets - issues Geoff Landis | [55] |
| Exoplanets Angelle Tanner | [56] |
To collect the complex information in an orderly manner, the following "work breakdown structure (WBS)" is proposed, along with an associated table of variables (table 6). By definition, a WBS is breakdown of a project into smaller, more manageable components, typically organized in a hierarchical structure. Detailed discussion of what these WBS elements mean and how they are measured is presented in the next two sections. The letters shown in parentheses are abbreviations for that WBS level, while the letters within brackets are the variables associated with that level (variables listed in table 6). The intent is to identify the least number of questions necessary to assess the
• Mission Choices, Top-Down o Destination (MD) [Dn, Dd, Di] o Mission Ambition (MA) [Wa] � Arrival Trajectory (MAT) [∆V, Do, Dr, Tot] � Science Sought (MAS) [Ir, Iv] o Timing (MT) [Ttp, Trl, Tdev (=Tscb+Tbd), Tm (=Tt+Ts+Tx)] o Motive (MM) [Wm] o Baseline Mission and Payload Scenarios (MB) � Solar Gravitational Lens � Deep Interstellar Medium � Centauri Flyby � Centauri Slower Flyby � Centauri Orbiter • Technical Challenges, Top-Down o Payload Challenges o Propulsion & Power Challenges o Incremental and Affordable Infrastructure Creation • Propulsion and Power Prospects, Bottom-Up o List of Propulsion and Power Prospects o Distinct Eras – Distinct Analyses � Era of Precursors � Era of Infrastructure � Era of Breakthroughs o Propulsion and Power Types (for energy calculations) (PP) � Type IP-OM: Internal Power & Onboard Reaction Mass � Type RP-OM: Received Power & Onboard Reaction Mass � Type RP-XM: Received Power & External Reaction Mass � Type IP-XM: Internal Power & External Reaction Mass o Infrastructure Dependence (I) � Earth-Based Beaming Infrastructure � Infrastructure Energy Availability & Usage Expense [Esi, Pia, Msc, Mb, Esc, Eb] � Mission Development Duration [Tdev] o Comparative Rates of Advancement (R) � Baseline Performance Trends of Shared Technologies & Resources � Technology Maturation Step Durations � Power & Propulsion Experiences Modeled as S-Curves • Flight Trajectory Analyses (FT) • Figures of Merit (FOM) o Mission Composite Value (maximize) [W] o Total Mission Expense (minimize)  [Etm (=Eti+Etp)] o Total Project Duration (minimize) [Ttp] o Mission Efficiency (maximize)  [We]
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Millis 2018 Grant NNX17AE81G_for_CR.docx The flow diagram in figure 5 is for guiding both the relative and deterministic analyses.
This version is expected to be refined during the Stage II work. For the relative analysis, this diagram represents a more detailed version of the topological analyses introduced with figure 1. It is expected that the analysis questions can flow more than one way. For example, one can either specify the total mission duration, data rate, and data volume to determine the required trip velocity, or specify the desired trip velocity and data rate to determine the total mission duration. The type of questions that the topological analysis aims to answer include:
• Which mission choices are the most difficult to achieve? • Which mission architectures have the greatest ratio of "composite mission value" to "total mission expense"? • Which mission architectures have the greatest ratio of "composite mission value" versus "total mission duration"? • Which technology elements within the propulsion and power options are more impactive to the final figures of merit?
This flow diagram is also to guide the more deterministic calculations to answer questions such as:
• Assess trades of "travel time" versus "data transmission time" within bounded "total mission durations." • Calculate how much energy it takes for the different propulsion and power options to reach 10%c. • Compare the payload mass delivery capability of different propulsion and power concepts as a function of trip time and destination distance.
The key differences between this analysis and prior interstellar mission analyses include: • Infrastructure dependence is included as part of the total mission cost and duration. • Different technologies are compared using the same mission and payload specifications. • The overall impact of different data transmission rates is part of the trade space. • Provisions are included for comparing the impact of different technology maturation rates.
Millis 2018 Grant NNX17AE81G_for_CR.docx

pg 29 of 69
•  Identify propulsion system type !  IP-OM !  SP-OM !  SP-XM !  IP-XM
•  Enter specs for that type
Proceed with [A] or [B] [A] Select mission–payload baseline [B] Input custom choices
∆V
•  Destination Dn, Dd, Di
•  Arrival trajectory, Do, Dr, Tot
•  Select 2 of 3 communication:
!  Data Rate, Ir !  Data Volume, Iv !  Transmission Duration, Tx
Total Propulsion
Spacecraft Launch Mass
•  Select 1 of 2 for journey time: !  Cruise speed, Vt !  Trip time, Tt
Energy
+
Etp (J)
Bese Mass (kg)
•  Launch Year (Yl, year)

|  |  |
| --- | --- |
| Infrast Depen | ructure dence |
| Infrastructure Use Specific Energy Esi (J/kg) Infrastructure Available Power Pia (W) Mission Development Duration, Tdev (yr) |  |

Total Infrastructure
Payload Mass, Mpay
Energy
•  Comm Sys SpPwr (W/bps)
Eti (J)
•  Payload Power (w)
•  Payload Spec Mass (kg/w) (Science Instru incl.)
Total Mission
KE Payload
Energy Etm (J)
Epy
(yr)
Total Mission Duration, Tmd (yr)
• Total mission expense, etm (J) • Total project duration, ttp, (yr) • Mission efficiency, we (%) • Mission composite value, W (#)
Total Project Duration, ttp (yr)

[Figure: Fig. 5. Basic Analysis Flow Diagram]

pg 30 of 69

Millis 2018 Grant NNX17AE81G_for_CR.docx are measured and their functional dependence are described in the next two sections.
m/s2
FT Acceleration a
I or O 8.4 FT Acceleration (average) ā
m/s2
I or O 8.4 FT Acceleration (time) Duration Ta s I or O 8.4 IP-XM FT Acceleration Distance Da m I or O 8.4 MTY Arrival Year Ya CY I or O 7.3 Goal or calculated MT Base Hardware Build (develop) Duration Tbd yrs O 8.3 Calculated from infrastructure J and W I Base Hardware Build Energy Eb J O 8.3 Calculated from infrastructure J/kg I Base Hardware Mass Mb kg O 8.3 Calculated from propulsion specs PPB Beam - Energy Cost of Beam Ebhw J or $ O 8.4 RP-OM, RP-XM PPB Beam Aperture  Area Ab
m2
I 8.4 RP-OM, RP-XM PPB Beam Director Cost $ I 8.4 RP-OM, RP-XM PPB Beam Director Cost per Aperture Area
$/m2
I 8.4 RP-OM, RP-XM PPB Beam Director Cost Scaling rad√$ I 8.4 RP-OM, RP-XM PPB Beam Divergence Angle Θ rad I 8.4 RP-OM, RP-XM PPB Beam Generating Power Pgb W I or O 8.4 RP-OM, RP-XM PPB Beam Power Generating Efficiency
% I 8.4 RP-OM, RP-XM PPB Beam Range Limit Dbl m I 8.4 RP-OM, RP-XM PPB Beam Range Limit Correction Factor % I 8.4 RP-OM, RP-XM PPB Beam Source Cost Scaling $/W I 8.4 RP-OM, RP-XM PPB Beam Transmitted Power Pbx W I or O 8.4 RP-OM, RP-XM MAS Closest Approach Do AU I 7.2 FT Coasting (time) Duration Tc s I or O 8.4 PPB Cost of Beam Source $ O 8.4 RP-OM, RP-XM FT Cruise Speed Vt m/s I or O 8.4 Goal or calculated MAS Data (Information) Rate Ir bps I or O 7.2 Corresponds to communication system MAS Data (Information) Volume Iv Mb I or O 7.2 Corresponds to communication system MT Data Transmission (time) Duration Tx s I or O 7.3 Corresponds to communication rate PPX Delta V of media by thrusting effect ∆Vt m/s 8.4 IP-XM MD Destination Distance Dd AU I 7.1 Pull from Table 1 MD Destination Level of Interest Di #rank I 7.1 Pull from Table 1 MD Destination Name Dn text I 7.1 Select from Table 1 PP Effective Exhaust Velocity (jet velocity) Ve m/s I 8.4 IP-OM, RP-OM R Energy Storage Energy Density f(y) J/kg I 8.5 Design standard trend to model R Energy Storage Power Density f(y) W/kg I 8.5 Design standard trend to model MTY First Data Arrival Year Yd CY I or O 7.3 I Infrastructure Available Energy Eia J I or O 8.3 Need to create per-decade estimates I Infrastructure Available Power Pia W/kg I or O 8.3 Need to create per-decade estimates I Infrastructure Use Specific Energy Esi J/kg I 8.3 Need to create per-decade estimates MAS Instrument Range Dr AU I 7.2 PP Jet Power Pj W I 8.4 IP-OM, RP-OM FOM Kinetic Energy of just Payload Epy J O 7.5 Total ∆V x Payload Mass MTY Launch Year Yl CY I or O 7.3 Goal or calculated PPX Mass Flow Rate of Media Thru Thruster dm/dt kg/s I 8.4 IP-XM MA Mission Ambition Wa # O 7.2
ηB
aV•∆V•aD•(Dr – Do)•aTTtot•aI•Iv
MTY Mission Commit Year Ym CY I or O 7.3 Actual starting point, assumes TRL >6 FOM Mission Composite Value W # O 7.5 Weighted sum of Di, Wa, Wm MT, I Mission Development Duration Tdev yrs O 7.3 Sum of Tbd and Tcsd FOM Mission Efficiency We % O 7.5 Epy/Etm MTY Mission End Year Ye CY I or O 7.3 When all data has reached Earth
Millis 2018 Grant NNX17AE81G_for_CR.docx

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PPB Particle Beam Mass Flow Rate dm/dt kg/s I 8.4 RP-XM PPB Particle Beam Velocity Vb m/s I 8.4 RP-XM PPB Particle Mass Expended Mp kg I 8.4 RP-XM MB, PP Payload Mass Mpay kg I 8.4 Baseline with respect to data rate PPB Pellet Specific Energy Esp J/kg I 8.4 RP-XM PPB Pellet Velocity Incident to Spacecraft Vip m/s I or O 8.4 IP-XM, RP-XM PP Propellant Mass Expended Mp kg O 8.4 IP-OM, RP-OM PP Propellant Mass Flow Rate dm/dt kg/s I 8.4 IP-OM, RP-OM PP Propellant Specific Energy Esp W/kg I 8.4 IP-OM R Propellant Specific Tankage Fraction f(y) Mt/Mp I 8.5 Design std trend to model (STS ET =3.7) PP Propulsion Power Source Specific Power 1/힪 W/kg I 8.4 IP-OM, RP-OM PP Propulsion System Specific Power Psp W/kg I 8.4 PP Propulsive Power P W I or O 8.4 R Radiator Specific Mass, f(K˚) f(y) kg/W I 8.4 Design standard trend to model PPB Receiver Aperture  Area Ab
m2
I 8.4 RP-OM, RP-XM PPB Receiver Areal Density
kg/m2
I 8.4 RP-OM, RP-XM PPB Receiver Flux Limit
W/m2
I 8.4 RP-OM, RP-XM MTY Research Commit Year Yrc CY I or O 7.3 MT Signal Time Ts yrs O 7.3 Function of Dd MT Spacecraft Build (develop) Duration Tscd yrs O 8.3 Calculated from infrastructure J and W I Spacecraft Build Energy Esc J O 8.3 Calculated from infrastructure J/kg PP Spacecraft Empty Mass (w/o Payload) Mse kg O 8.4 Calculated from propulsion specs PP Spacecraft Launch Mass Msl kg O 8.4 IP-OM version: Mse+Mpay+Mp PP Specific Impulse Isp s I or O 8.4 MM Sum of Motives Wm # O 7.4 (needs graduated scale) R Tech Maturation Increment Duration f(Li) ∆Li yr O 8.5 Calculated (TBD) R Tech Maturation Increment Pattern  f(Li) ∆Li yr I 8.5 Historic relative duration between steps R Tech Maturation Level (TRL, SML) L # I 8.5 Table 13 MT, R Technology Maturation Duration Trl yrs O 7.3 Time between now and TRL 6 PP, FT Thrust F N I 8.4 PP Thrust Conversion Efficiency η % I 8.4 MA Time on Target Tot sec I or O 7.2 Time that spacecraft within instru' range
∆V m/s I
7.2
The total velocity change for mission FOM, I Total Infrastructure Energy Expense Eti J O 8.3 Eti = Es + Eb MT Total Mission Duration Tm yrs I or O 7.3 Sum of Tt + Ts + Tx FOM Total Mission Expense (energy) Etm J O 7.5 Sum of Etp + Eti FOM, MT Total Project Duration Ttp yrs I or O 7.5 Time between Yrc and completion (Ym) FOM Total Propulsion Energy Etp J O 8.4 Energy for all mission propulsion MT Trip Time Tt yrs I or O 7.3 FT Velocity of Spacecraft Vs m/s I or O 8.4 RP-XM PPX, FT Velocity of Spacecraft Through Media Vsm m/s I or O 8.4 IP-XM
pg 32 of 69

Millis 2018 Grant NNX17AE81G_for_CR.docx The starting point for a mission is deciding where to go, what to do once there, and how soon to get there.
The first question is, "Where to go?" Already there is trade between easier missions and more interesting destinations. Typically, the more interesting destinations are further away. The variables for this part of the analysis are:
Destination Name, Dn: The name of the chosen destination as selected from table 1, which then has corresponding other measures; distance and level of interest.
Destination Distance, Dd, (AU): The distance to the destination, measured in AU or ly. For most cases this is a simple number, but there are destinations that span large distances (such as the Oort cloud), plus the special case of the "Oumuamua" object, which is a fast moving extrasolar object (5 AU/yr) that is still relatively close (2018, 3 AU) [78].
Destination Interest, Di: This is a subjective value, where different people might have different notions of what is more or less interesting. For the sake of this study those subjective differences are cast into specific values, which can later be debated if found that the choices heavily sway technical priorities. Those provisional values are listed in the "Level of Interest" column of table 1.
Almost all interstellar mission concepts assume a flyby at whatever coast velocity the spacecraft has achieved, and where the closest approach is determined by how accurately the spacecraft was originally aimed. Anything else requires additional propulsion and the associated increases in mass and power (and subsequent increases in trip time or required energy). To more systematically include the trajectory options in the analyses, the following categories of arrival trajectories are listed in order of increasing ∆V in table 7:

| Arrival Trajectory Options | Initial Acceleration | Course Correction | Braking | Orbit Insertion |
| --- | --- | --- | --- | --- |
| Flyby Fast (cruise velocity) | Maximum Possible | Maybe | 0 | 0 |
| Flyby Slower | Maximum Possible | Maybe | Some | 0 |
| Orbit Star | Maximum Possible | Yes | Significant | Yes |
| Orbit Exoplanet | Maximum Possible | Yes | Significant | Yes |

Total ∆V (m/s): Calculated as usual. In principle, the sum of the individual ∆V's in a row in table 7.
Millis 2018 Grant NNX17AE81G_for_CR.docx

pg 33 of 69

Closest Approach, Do (AU): Presumably measured in AU, this is defined by how close the probe passes the target star or exoplanet. Closest approach is affected by the initial aiming accuracy of the spacecraft, plus any maneuvering capability later. The maneuvering capacity is already measured as a part of the total ∆v.
Instrument Range, Dr (AU): This value is a function of the technology used for the scientific instruments. In the absence of existing specifications, fixed test-case values will be assigned.
Time on Target, Tot (s): Presumably measured in hours or days, time on target is how long the probe is within the instrument range. For flyby missions, this is inversely proportional to the cruise velocity. For orbital missions it is equal to the remaining functional life of the spacecraft. Table 8 shows different times-on-target for six different instrument ranges and five different flyby speeds. For example at 20% lightspeed, a spacecraft would only be within ± 1 AU for less than 90 minutes. Even relaxing that to only 1% lightspeed and a range of ± 5 AU, the spacecraft will only be in that range for 139 hours.
As a representative example of reasonable minimum viewing durations, it is fitting to examine how much time it takes for an Earth-viewing satellite to integrate enough images to be able to subtract the cloud interfearance and reveal the ground features. Is this duration hours, days, or months?

| Flyby Speed | Instrumentation Range ± AU 0.5 1 5 10 50 100 |  |
| --- | --- | --- |
| (c) 0.01 0.05 0.10 0.15 0.20 | ––– hours ––– 14 28 139 2.8 5.5 28 1.4 2.8 14 0.9 1.8 9.2 0.7 1.4 6.8 | ––– days ––– 12 58 116 2.3 12 23 1.2 5.8 12 0.8 3.8 7.6 0.6 2.8 5.7 |
What scientific instruments should an interstellar probe carry to collect information that cannot be obtained from Earth-based astronomy alone (even after decades of further advances in astronomy)? Would the trajectory's closest approach and time on target be sufficient to collect this information and how much data is sufficient to reach meaningful conclusions?
NOTE: Before proceeding to quantify the science ambitions, this topic presents an excellent example of the utility of precursor missions to this study. As stated before, precursor missions are not assessed as part of this study, but do provide suitable baseline examples. To help resolve the question of required instrumentation, range, and time on target, a fitting precursor mission would be a "Look Back Mission." A suite of exoplanet instruments can be tested by looking back toward Earth at various distances (1-100 AU?) to determine the required closest approach and time on target for collecting meaningful information.
The basic trade here is between smaller, less capable payloads that can reach the destination sooner, versus larger, more capable payloads that will take longer to get there. As discussed previously, the total mission duration includes the time to transmit the data, which is a function of the amount of data taken and the data transmission rate.
Data Rate, Ir (bps): The trades here involve the choice between sending a small and fast payload with long data transmit time, or larger, slower payload with shorter transmit times.
For example, the StarShot mission allocates 20 years for transmitting the data back to Earth, where that very long duration enables the use of an incredibly low power and low mass communication system on the payload. Since this is a significant proportion of total mission duration, it is an important factor to specify. Data return time is a function of the data rate, measured in bits per second (bps), and the total amount of data, measured in bits (or bytes, 8 bits = 1 Byte).
Data Volume, Iv (bits or Bytes): How much date is enough data? Like the question of what will be measured, there is no consensus on the total quantity of data. Since greater quantities of data can extend the total mission duration, the total quantity of data is considered in the trade space.
pg 34 of 69

Millis 2018 Grant NNX17AE81G_for_CR.docx As examples of what might be considered sufficient data, consider some of the first historic missions that imaged the Moon and Mars, such as Luna 3 (first mission to image the Lunar far side)

and Mariner 4 (first Mars flyby) both of which returned roughly 8 Megabits of data (1 MB). Here are examples of images from those first flyby missions.
Composite image of the backside of the Moon
Image of Mars Source: https://nssdc.gsfc.nasa.gov/imgcat/hires/lu3_2.gif

[Figure: Fig. 6. "First Image" Examples from History]

Estimated to be equivalent to 8 Megabits (1 MB) each

Extending the analogy to a possible orbiter mission, consider the Mariner 9 mission, the first Mars orbiter. That mission produced roughly 16 Gigabits (2 GB) of data. While these data quantities are small compared to recent planetary probes, such as Cassini-Huygens at over 600 GB, that was a Saturn probe at only about 10 AU from Earth. That's 27 thousand times closer to Earth with communication that much easier. For our first interstellar missions, what are the lower bounds of what would be acceptable images? For example, a single image of an exoplanet in 6 bit gray scale, with 30 pixels across the equator (which would represent a tremendous leap beyond anything currently possible with astronomy) is about 4200 bits (0.5 kB).
Data Fidelity, If (?): Perhaps an additional factor could be how much is learned from a given volume of data, though how to quantify such a parameter is uncertain.
Mission Ambition," Wa (#): Finally, these factors combine to give a relative score for Mission Ambition. Since the individual terms use different units, normalizing and weighting coefficients, a, must be included for each term. A tentative equation for Mission Ambition is:
Eq. 2

) aTTot aI IV Eq. 2
Wa = aV ΔV aD Dr −Do where higher interest is proportional to more time on target, higher data return, useful science at more distant instrument ranges, and more capable propulsion systems denoted by higher ΔV capabilities. Note that if the closest approach, Do, is greater than the instrument range, Dr, then the score for Mission Ambition goes negative. In other words, if the spacecraft cannot get close enough to the destination for the science instruments to do their job, then the mission is a failure. Additional variations of Mission Ambition can be considered in developing an optimized metric.
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