FISSION FRAGMENT ROCKET: Fuel Production and Structural Considerations Pauli Erik Laine FLYING on a RAINBOW a Solar-Driven Diffractive Sailcraft Grover A

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Journal of the British Interplanetary Society

VOLUME 71 NO.4 APRIL 2018 Tennessee Valley Interstellar Workshop

PULSED MAGNETIC NOZZLE for Fusion Propulsion Jason Cassibry et al FISSION FRAGMENT ROCKET: Fuel Production and Structural Considerations Pauli Erik Laine FLYING ON A RAINBOW A Solar-Driven Diffractive Sailcraft Grover A. Swartzlander, Jr. EVALUATION OF THE HAZARD OF DUST IMPACTS on Interstellar Spacecraft Richard A. London & James T. Early A SCIENCE-DRIVEN MISSION CONCEPT to an Exoplanet Stacy Weinstein-Weiss et al CONTACT WITH ALIEN BIOMES: Possible Biochemical Incompatibilities Kenneth Roy & Catherine Smith

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118 INTRODUCTION Tracie Prater and Les Johnson

119 PULSED MAGNETIC NOZZLE for Fusion Propulsion Jason Cassibry et al

126 FISSION FRAGMENT ROCKET: Fuel Production and Structural Considerations Pauli Erik Laine

130 FLYING ON A RAINBOW A Solar-Driven Diffractive Sailcraft Grover A. Swartzlander, Jr.

133 EVALUATION OF THE HAZARD OF DUST IMPACTS on Interstellar Spacecraft Richard A. London & James T. Early

140 A SCIENCE-DRIVEN MISSION CONCEPT to an Exoplanet Stacy Weinstein-Weiss et al

150 CONTACT WITH ALIEN BIOMES: Possible Biochemical Incompatibilities Kenneth Roy & Catherine Smith

OUR MISSION STATEMENT The British Interplanetary Society promotes the exploration and use of space for the benefit of humanity, connecting people to create, educate and inspire, and advance knowledge in all aspects of astronautics.

JBIS Vol 71 No.4 April 2018 117 INTRODUCTION

Introduction by TRACIE PRATER1 and LES JOHNSON2, 1Materials scientist at NASA’s George C. Marshall Flight Center and presenter at the TVIW; 2General Chair of the TVIW, and Principal Investigator of NASA’s first interplanetary solar sail mission, Near Earth Asteroid Scout.

“All civilizations become either spacefaring or extinct.” – CARL SAGAN

f all the journeys the human species has imagined plasma environment. Other papers consider laser-driven light- in the course of its 100,000 year history, none is sails as a proposed propulsive technique for interstellar travel. longer, more daunting, and more fraught with In “Flying on a Rainbow: A Solar Driven Diffractive Sailcraft”, technical and physiological challenges than the Grover Swartzlander offers an alternative to the traditional sail Ojourney to the stars. From our earth-based per- design, which uses highly reflective materials, by proposing an spective, interstellar space begins where the influence of our active diffractive sail surface. Absorption is significantly -re Sun ends. The boundary is marked by a sharp increase in the duced and transmitted photons can be re-used or exploited to density of plasma far beyond levels observed in the heliosphere provide energy for craft subsystems. (the region where the solar wind – the flow of charged particles from the sun – is present) and the heliosheath (the outer- Regardless of the propulsion technique used, an interstel- most shell of charged particles emanating from the sun). This lar craft must be able to survive the harsh environment of the space between stars has to date been traversed only by a single interstellar medium. Richard London’s “Evaluation of the Haz- human-made probe, Voyager 1, which left the heliosheath in ard of Dust Impacts on Interstellar Spacecraft” examines the 2012 after 36 years of operation. The significance of Voyager 1 impact of space environmental effects on a craft traveling at is immense – quite simply, it represents humanity’s first, small a fraction of the speed of light in the region beyond the heli- step into the interstellar medium. Moving at 17 km/s, the probe osheath. London simulates the collision of energetic interstellar continues its journey at only 0.006 percent of the speed of light. dust grains with structures of varying thickness and, based on these analyses, offers design strategies to mitigate damage. Held 40 years after the launch of Voyager 1, the October 2017 meeting of the Tennessee Valley Interstellar Workshop Another paper presents system-level analyses of interstellar (TVIW), with its theme of “Step by Step: Building a Ladder to missions. Stacy Weinstein-Weiss’s “A Science Driven Mission the Stars”, emphasized the small steps that will one day pave Concept to an Exoplanet” designs an interstellar mission with the way for monumental interstellar missions. TVIW is a non- science requirements as the preeminent consideration. In mis- profit organization with a goal of facilitating research, explora- sion formulation, the authors trade propulsion, power, commu- tion, and education in the field of interstellar travel. The papers nications, instrumentation, and navigation systems to enable an collected here originate from the TVIW 2017 conference and interstellar mission with the objective of life detection. And the address the litany of challenges humanity will face as we seek to search for life on other worlds is addressed by Ken Roy in the travel farther and faster through our cosmos. “How will we get paper “Contact With Alien Biomes: Possible Biochemical In- there?” is the fundamental question at the center of many tech- compatibilities,” which addresses the potential incompatibilities nical conversations related to interstellar travel. If humans are of human biology in extraterrestrial environments. to ever undertake interstellar missions or data from interstellar probes is to be returned during the lifetime of a scientist, transit The Journal of the British Interplanetary Society (JBIS) has speeds for spacecraft must be drastically increased. a rich and storied history of publishing original research in the field of interstellar travel. Many of the landmark publications in The selected papers from TVIW present technical analyses the field first found a home in this journal. We believe the arti- of several propulsion concepts to enable interstellar journeys. cles represented in this issue contribute to that legacy and hope Jason Cassibry’s “Pulsed Magnetic Nozzle for Fusion Propul- they will facilitate renewed conversation and new research on sion” considers nozzles for fusion or pulsed fission/fusion hy- maturing the technologies needed to undertake this grandest of brid propulsion systems. In this design, which stands in contrast human challenges. The “why” of interstellar travel is directly ad- to the traditional pusher plate and solid state nozzle concepts, dressed in many of these papers, but perhaps was also succinctly plasma does not come in direct contact with the coils, presum- captured in the words of John F. Kennedy in his seminal speech ably leading to less erosion and a longer operational life. Pauli on the US space program at Rice University in 1962: “We set sail Erik Laine discusses the design challenges of a craft propelled on this new sea because there is new knowledge to be gained…. by a fission reaction in “Fission Fragment Rocket: Fuel Produc- its conquest deserves the best of all mankind, and its opportuni- tion and Structural Considerations.” Laine focuses on the limi- ty for peaceful collaboration may never come again.” tations in production rates for the nuclear isotopes used as fuel, safety concerns with fission systems launched from earth, and Ad astra, the associated structural and thermal design considerations of a spacecraft moving at interstellar speeds through a dense Tracie Prater, Les Johnson

118 Vol 71 No.4 April 2018 JBIS JBIS VOLUME 71 2018 PAGES 119-125

PULSED MAGNETIC NOZZLE for Fusion Propulsion JASON CASSIBRY, BRYAN WINTERLING AND KEVIN SCHILLO, Propulsion Research Center, University of Alabama in Huntsville, Huntsville, AL 35899 USA email [email protected]

A fusion or pulsed fission/fusion hybrid propulsion system has the potential to enable rapid interplanetary, deep-space and interstellar precursor missions. For these systems, one of the problems to be solved is the design and development of a nozzle to convert isotropic thermal expansion of a burning target into directed thrust. This paper presents three dimensional simulations of magnetic nozzles to generate propulsive thrust from deuterium tritium plasma. Of interest, was the discovery of a novel winding of field coils, longitudinally rather than azimuthally, to help promote a higher component of jxB Lorentz force in the axial direction. Gravitational lensing and interstellar precursor missions were investigated using gravity free straight line trajectory analysis to determine notional performance requirements to assess the needed improvements in the approach and results presented in this study.

Keywords: Fusion propulsion, Fission/fusion hybrid propulsion, Magnetic nozzle

1 INTRODUCTION

A fission/fusion hybrid or pure fusion propulsion system offers a potential path to enabling rapid interplanetary travel. Using straight line trajectory mission analysis assuming continuous thrust for 2/3 of the trip distance (1/3 acceleration, 1/3 coast, 1/3 deceleration each way), round trip times for notional deep- space sample return missions using a vehicle with a specific power of 1 kw/kg are plotted in Fig. 1 for destinations from Jupiter out to the dwarf planet Eris. The travel time to and from Jupiter can be less than three years, while samples return from the dwarf planets out to Eris can be accomplished between 10 and 17 years. Studies conducted by Cassibry et al [1], Miernik round trip time (years) et al [2] and Adams et al [3] have also shown that round trip, human piloted missions to Mars can be two to six months: thus, distance (AU) greatly reducing the need to wait for period launch windows. Although more demanding, interstellar precursor missions have also been considered and are summarized in an extensive Fig.1 Comparison of roundtrip times for a 1 kW/kg fusion report by Schulze [4], with flyby times ranging from ~50 years propulsion powered spacecraft to destinations spanning from to a few centuries depending on assumptions about propulsion Jupiter to Eris. The locations of the outer dwarf planets are technology and destination. estimated based on anticipated distance from a 2017 launch date from LEO. A fusion propulsion system must have a mechanism to redirect an expanding fusion plasma in an axial direction in this plate, however in an actual system, ablation of the pusher order to impart a propulsive momentum on the spacecraft. A plate may contribute a non-negligible thrust component which simple option, explored during Project Orion in the 1950s and would require experiments to quantify. A slight variant investi- 60s, suggested a pusher plate at the rear of the spacecraft. Nu- gated by the authors of this paper involved replacing the pusher clear explosives were to be ejected and detonated at a certain plate with a hemispherical nozzle, which was found to offer a distance from the spacecraft, with the plasma generated by higher specific impulse [6]. There have been some analytical the explosion then impacting the pusher plate and propelling models and engineering design of coils, such as in the Human the vehicle forward [5]. The primary momentum imparted is Outer Planet Exploration (HOPE) study [3]. The most recent due to the stagnation and expansion of the propellant against and rigorous work in this area was a 2D MHD simulation of a parabolic nozzle shape, building on the HOPE study [7]. These preliminary results suggest that plasma reflection and This paper was presented at the Tennessee Valley Interstellar redirection can be achieved but instabilities at the vacuum/ Workshop 2017 Symposium, Huntsville, USA. plasma boundary interface introduce difficulties which must

JBIS Vol 71 No.4 April 2018 119 JASON CASSIBRY et al be addressed. This paper presents numerical results for a 3D pulsed magnetic nozzle, which may offer an improved propul- (5) sion mechanism over a pusher plate or solid-state nozzle, with anticipated longer lifetimes since the plasma does not come in direct contact with the coils generating the nozzle. To the authors’ knowledge this is the first such work in a 3D simulation. and the first moment is:

2 NUMERICAL METHOD (6) All simulations are performed with the Smoothed Particle Fluid with Maxwell equation solver (SPFMax) [8]. Smoothed Particle Hydrodynamics (SPH) [9] is a meshless Lagrangian method that simulates hydrodynamics by dividing a fluid into Higher order moments such as second order derivatives a set of particles and using a summation interpolant function also sum to 0. The particle volume is: to calculate the properties and gradients for each of these particles. In simple terms, the SPH method is nothing more than an algorithm for interpolation of properties and gradients from (7) scattered data in 1, 2, and 3 dimensions. There are two approximations that are made. First, the interpolation of a property of a field can be determined by the integral approximation and η is 1.11 in SPFMax, and variations on that value will require adjustments to h so that the constraints remain satisfied. (1) Thus, the gradients in a fluid or plasma are calculated at a specific particle location using the particle volume and proximity of other nearby particles [10] [11]. For second order derivatives where A is any property (like temperature, pressure, etc), sub- in SPFMax, the code stores first order derivatives as particle script a means point or particle a, r is the position of point properties in which second order derivatives are needed [12], a in space, and W is the interpolating kernel function. In the [13]. This includes the three spatial derivatives of each of the limit that h→ 0, W becomes the direct delta function and the three velocity components [14, 15] as well as the scalar temper- expression becomes exact. Any function can be approximated ature for the heat conduction equation [16]. Details of the SPH in this way, and this is the first assumption of SPH. The second method can be found in Ref. [9]. assumption is to replace the integral with a summation, SPFMax solves conservation of mass exactly because the continuity equation is not necessary. Rather, density is de- (2) termined by the particle mass divide by the particle volume, where the mass is a constant property of the particles. The momentum equation for a single fluid is given by: where Vb is the volume of the neighboring particles b. This is called the summation or particle approximation, and is anal- ogous to discretizing a computational volume into a mesh for (8) finite difference or finite element algorithms. The kernel -func tion W is usually a Gaussian-like or cubic b-spline function which goes to zero at some κh, where κ=2 normally. Gradients where p is the static pressure and τ is the deviatoric viscous can be approximated as stress tensor. The single temperature energy equation (Ti = Te = T) is given by:

(3)

In SPFMax, the cubic spline function is used, (9)

where k is the thermal conductivity, σ is the Stefan-Boltz- mann constant, T is temperature and χPlanck is the single group Planck emission opacity. Alternatively, if the optical thickness 1⁄(ρχ Planck) is of the same order or smaller than the particle scale h, then radiation can be modeled as a diffusion process. This is achievable by adding an additional term to the overall thermal conductivity [17]: (4) where q is the ratio of the radius from particle a to the neigh- bor particle b to the particle h size. The key to implementing (10) any SPH method properly is to have an accurate list of neigh- bors for each particle and a compact support distance h which scales the kernel function and its gradients so the following constraints are satisfied: where a=4 σ⁄c is the radiation density constant.

120 Vol 71 No.4 April 2018 JBIS PULSED MAGNETIC NOZZLE for Fusion Propulsion

If the temperature is split between ions and electrons, then the two temperature energy equations are given by: (a) (b)

(11) and

(12) (c) where

(13)

and the electron collision time is[25] Fig.2 Magnetic nozzle shapes that were attempted but proved unsuccessful include (a) NERVA-shaped solenoid, (b) single-turn and (c) bell-shaped nozzles. (14)

4 RESULTS

In SPFMax, several magnetic field solvers have been ex- Figure 2 shows three of the initial coil geometries and magnetic plored. This has been a challenging part during the develop- field topologies attempted prior to arriving at a more successful ment of this code. The challenge lies within the desire to have a design. The failed ideas are presented here in hopes that others model that handled self-consistently external circuits coupled can make them function or independently verify the potential to the computational domain as well as electric and magnetic of these designs. Since there has been no experimental or 3D field propagation across vacuum, dielectric and conductors. simulation results on pulsed magnetic nozzles, the first attempt The basic approach [8] is that the plasma conducts current as was a solenoidal field coil where the windings were in the shape a network of transmission lines integrating Kirchoff’s voltage of and to scale with the NERVA thermal rocket engine (Fig- and current laws in which additional physics include back ure 2a). This shape was chosen because a starting point was electromotive force (EMF), the Hall Effect and etc. as needed. needed and NERVA was an experimental program in advanced Field propagation is accomplished by utilizing a superposition propulsion with hard data. Two types of simulations were at- of current sources using Biot Savart’s law. It was found that this tempted: one with the plasma born on a steady state field and brute force method was the simplest and least error prone way one in which the coil was pulsed. In the former, an initial ve- to enforce divergence of B equaling 0. locity of 20 km/s was applied; as a result, with pre-magnetiza- tion, the plasma was accelerated back towards the wrong end of 3 APPROACH the nozzle. This was similar to a refrigerator magnetic resisting the pull of removal from a ferromagnetic surface. In the pulsed Simulations were conducted to explore different magnetic configuration, a surface current was induced on the plasma topologies resulting from variations in solenoidal windings. inside the nozzle and some minor acceleration was observed; The three winding variations explored were ring shaped, bell however, the primary force was in the radial direction. In such shaped and a winding in the shape of the nozzle developed in cases, ranges of current, pulse width and etc. were all tested the 1960s for nuclear thermal propulsion, The Nuclear Engine and tried. It is important to note that the ratio of the magnetic for Rocket Vehicle Application (NERVA). A major parame- field pressure to ram pressure during thermal expansion must ter explored in these simulations found a magnetic topology exceed unity in order for the nozzle to work. Therefore, it was that provides a Lorentz force in the axial direction as the fu- determined that a nozzle shape with a larger component of ra- sion plasma expands. A deuterium tritium plasma target was dial field was needed due to minimal acceleration observed in introduced inside each coil geometry and allowed to expand the NERVA-shape. thermally with an initial density of 80 kg/m3 and temperature of 1 keV. The target was 1 mm in radius, 1 cm in length and had Figure 2b shows the second nozzle variation, a single turn a mass of 2.5 mg. The reason for the choice of density is due ring, in which a larger component of radial field is visible from to the stopping power of the 4He, which gives a range compa- the stream lines. Nevertheless, there was no significant accel- rable to the target radius at these densities. This is important eration because the field lines in the radial direction are gener- for achieving a high fractional burnup in the target, which is ally weak in comparison with the axial field inside the ring. In assumed to produce an average thermal temperature of 1 keV Figure 2c, a bell shape, like the bell of a trumpet, was devised in the fuel and propellant. It is clear that these assumptions as to once again to increase the strength of the radial field; only well as the feasibility of how to achieve these high densities in a modicum of improvement was observed. In all three cases a DT plasma experiment will have to be explored much further it was noted that the plasma experiences much stronger com- beyond this initial study. pression/confinement than expansion. The field strength under

JBIS Vol 71 No.4 April 2018 121 JASON CASSIBRY et al

(a) (b)

Fig.3 Magnetic nozzle shape that achieved a meaningful specific impulse, a) coil winding, b) magnetic field stream lines (nearly azimuthal) and contour flood plot of magnetic field strength, c) magnetic field pressure in [Pa] and d) nozzle with ablator plate (green) and initial target (magenta). the coils was always stronger in the axial direction, so the j×B that directs the flow of electrons to a phosphor screen for dis- Lorentz force is pointed radially inward. To achieve a larger playing an image. In this winding, the magnetic field strength component in the axial direction, fundamental departure from and direction are shown in Figure 3b. Noticeably, the field is al- the solenoid-family of nozzle shapes was taken. most 100% azimuthal in this geometry and the strength of the field is nearly 0 along the axis. The field strength grows rapidly It was hypothesized that it may be easier to generate strong in the direction of the field coils. The strength of the field in azimuthal fields and rely on thermal expansion to induce ra- effect takes on the shape of a nozzle. In Figure 3c, the magnetic dial currents; thus, increasing the force component in the ax- field pressure is shown with slices in the x=0 and y=0 planes; ial direction needed for thrust (Fig. 3). The winding can be in one case, it is in plane with two opposing wires, while in thought of as a central conductor through which current flows. the other plane, it is in between pairs of wires, thus indicating Ultimately, near the plasma target, the current is split into a that the field pressure is fairly uniform. In Figure 3d, an ablator number of individual wires that flare in the shape of a nozzle plate and initial target are shown to indicate the initial position with radial and axial components. At the nozzle exit, the wires of the plasma. are bent and return at some fixed radius. The wires will bend again to close the circuit back to the common center conductor. Figure 4 shows the initial position of the plasma and abla- The winding in this geometry was subsequently recognized to tor plate to scale at t=0 and at 80 ns. The plasma expands at resemble a deflection yoke, which was the technology used in about 350 km/s, comparable to the initial sound speed. At 80 cathode ray tubes in older televisions and analog oscilloscopes ns, the surface reaches the perimeter of the ablator plate. A few

(a) (b)

Fig.4 Expansion of cylindrical target with ablator plate shown for reference. In a), the target is shown to scale at t=0. In b), the time is 80 ns, the surface represents the boundary of the thermally expanded plasma, and vectors are visible indicating the production of current at the surface due to the back emf (expanding plasma against a steady field).

122 Vol 71 No.4 April 2018 JBIS PULSED MAGNETIC NOZZLE for Fusion Propulsion particles penetrate the plate, which was an error that was subsequently corrected by increasing the particle resolution in the plate. The radial expansion of the plasma moves into a rapidly changing magnetic field, which creates a motional back EMF that induces an image current in the plasma. This can be seen with the vectors that are mostly near the surface of the expanded plasma--in the direction antiparallel to the direction of the current in the coils. It should be noted that the case crashed at about 250 ns. Reasons for this are being explored and will be reported as this research matures.

The case that crashed was rerun without any field to try and isolate the ablator contribution to the thrust and specific impulse with that of the field. Figure 5 gives a time history of the axial (positive z-direction component) specific impulse for the target. The ablator plate with no field gives a peak 5000 seconds, while the magnetic nozzle reaches about 9000 seconds. This Fig.5 Comparison of ablator plate and magnetic nozzle specific case is in no way optimized; however, the fact that the plasma impulse versus time. remained stable and the magnetic nozzle provided marked improvement to the specific impulse achieved was encouraging. extra-solar planet and thus would serve as a precursor for an interstellar mission. The second potential mission was the flyby With confidence in at least one path to a working magnetic of Alpha Centauri. To perform the calculations, a straight line nozzle, two interstellar-relevant deep-space missions were ana- trajectory analysis was performed using the ‘Type II’ perfor- lyzed to determine the required propulsion performance. The mance equations developed and summarized by Moeckel [18]. first potential mission--made possible by a fusion propulsion Fig. 5 gives the mission performance for the parametric anal- system—was the deployment of a space telescope at a gravita- ysis in which trip time was varied. The payload was assumed tional lensing point, beginning at a distance of 550 AU. Such to be 10 metric tons with a propulsion system specific power a telescope would be capable of providing direct images of an of 10 kW/kg.

(a) (b)

Fig.6 Parametric analysis for fusion propulsion flyby missions to a gravitational lensing focal point (550 AU) and Interstellar precursor (269,000 AU, the approximate distance to the Alpha Centauri system).

JBIS Vol 71 No.4 April 2018 123 JASON CASSIBRY et al

Fig.7 Propellant mass flow rate and pulse frequency required for a pulsed fusion propulsion system for the parametric mission analysis presented in Figure 6.

Additionally, the mass flow rate was derived from the thrust sion or other pulsed systems, will be attainable with improved and specific impulse, and taking the initial target mass from nozzle design, increased target temperature with due consid- the simulations, the required pulse frequency was determined, eration to managing radiation loss and finite burn during ex- Figure 7. The mass flow rate is of the order of milligrams per pansion to maintain high pressures in the target in order to second, and the required pulse frequency was of the order of augment the rate of expansion. 1 to 10 pulses per second, although results could vary at the very low and very high trip times in which mass is traded for 5 CONCLUSIONS mission trip time. 3D simulations of a magnetic nozzle were performed using a Of interest is the selection of a vehicle mass at the knee of SPH code coupled with SPFMax. Results suggest that magnetic the curve; thus, trading trip time for smaller vehicle masses. nozzles are indeed capable of redirecting thrust from a plasma Table 1 provides the performance parameters needed to enable and found that coil shapes that create a mostly azimuthal mag- a mission to a gravitational lensing point as well as to Alpha netic field appear to produce a promising path toward efficient Centauri. A vehicle mass of 15.3 metric tons was picked for coil design. A specific impulse of 9000 s was achieved in such the gravitational lensing missions, which gave a trip time of a nozzle with a cylindrical deuterium tritium plasma at 1 keV 10 years. The final burnout velocity was 392 km/s and requires temperature, and the design was far from optimized. A null a specific impulse of 94,000 seconds. For the interstellar flyby case was run without the field turned on, achieving only half mission, the vehicle mass was selected at 1000 tons to bring of this specific impulse from the pusher plate, which was pres- down the lengthy trip time to 269 years. The specific impulse ent in both cases. Mission analysis was performed to guide the required for this choice of mass was 270,000 seconds, which needed improvements in the overall design, coupling fusion approaches the upper limit achievable by thermal expansion of to magnetic nozzle performance. A specific impulse of 94,000 a fusion plasma. For both missions, the thrust is of the order of seconds is required for a 10 year flyby of a 550 AU gravitational 1 N. Furthermore, the pulse frequency needed to achieve this lensing mission, while 270,000 s Isp is needed for a 269 year fly- varied from about three shots per second for the gravitation- by of Alpha Centauri. Future work will include refinements to al lensing mission to nine shots per second for the interstellar the coil design and burn physics during the thermal expansion mission. Such pulse frequencies are readily achievable in exist- to increase the efficiency and energy added to the plasma. It ing pulse power systems, such as the capacitor bank at the Nike appears that these high specific impulse requirements are pos- Laser Facility. sible to enable such missions.

Comparable to Table 1, the specific impulse produced from Acknowledgments the simulation is about a factor of 10 too low; however, the efficiency was about 20%. The route to the requisite specific im- This work was supported in part by NASA contract NN- pulse to make such deep-space missions achievable, with fu- M11AA01A.

TABLE 1 Input parameters of the simulation Trip time Isp IMLEO Thrust Shot frequency Mass flow rate ΔV Destination (years) (s) (metric tons) (N) (1/s) (mg/s) (km/s)

Gravitational Lensing 10 9.4x104 15.3 0.66 2.9 0.72 392

Alpha Centauri 269 2.7x105 1000 1.39 8.8 0.48 7,280

124 Vol 71 No.4 April 2018 JBIS PULSED MAGNETIC NOZZLE for Fusion Propulsion

REFERENCES

1. J. Cassibry, R. Cortez, M. Stanic, A. Watts, W. Seidler, R. Adams, G. 13. M. A. Rodriguez and J. T. Cassibry, "A 3-D Smoothed-Particle Statham and L. Fabisinski, "Case and Development Path for Fusion Hydrodynamics Model of Electrode Erosion," IEEE Transactions on Propulsion," Journal of Spacecraft and Rockets, vol. 52, no. 2, pp. 595- Plasma Science, vol. 45, no. 11, pp. 3030-3037, 2017. 611, 2015. 14. J. Bonet and T. S. L. Lok, "Variational and momentum preservation 2. J. Miernik, G. Statham, L. Fabisinski, C. Maples, R. Adams, T. Polsgrove, aspects of Smooth Particle Hydrodynamics Formulations," Computer S. Fincher, J. Cassibry, R. Cortez, M. Turner and T. Percy, "Z-Pinch Methods in Applied Mechanics and Engineering, vol. 180, pp. 97-115, fusion-based nuclear propulsion," Acta Astronautica, vol. 82, pp. 173- 1996. 182, 2013. 15. S. J. Watkins, A. S. Bhattal, N. Francis, J. A. Turner and A. P. Whitworth, 3. R. Adams, R. Alexander, J. Chapman, J. Fincher, S. Philips, A. Polsgrove, "A new prescription for viscosity in Smoothed Particle Hydrodynamics," T. Wayne, B. Patton, G. Statham, S. White and Y. Thio, "Conceptual Astronomy and Astrophysics, vol. 119, pp. 177-187, 1996. Design of In-Space Vehicles for Human Exploration of the Outer 16. J. H. Jeong, M. S. Jhon, J. S. Halow and J. v. Osdol, "Smoothed particle Planets," NASA Technical Rept. 2003-212691, 2003. hydrodynamics: Applications to heat conduction," Computer Physics 4. N. Schulze, "Fusion Energy for Space Missions in the 21st Century," Communications, vol. 153, no. 1, pp. 71-84, 2003. NASA Office of Safety and Mission Quality, NASA Technical 17. Y. B. Zel'dovich and Y. P. Raizer, Physics of Shock Waves and High- Memorandum TM4298. Temperature Hydrodynamic Phenomena, Meneola, New York: Dover 5. "Nuclear Pulse Space Vehicle Study," George C. Marshall Space Flight Publications, Inc., 2000. Center, Huntsville, AL, 1964. 18. W. E. Moeckel, "Comparison of Advanced Propulsion Concepts for 6. K. Schillo, Three-Dimesional Modeling of an Ideal Nozzle for Addvanced Deep Space Exploration," Journal of Spacecraft and Rockets, vol. 9, no. Propulsion, a thesis, Huntsville, AL: University of Alabama in 12, pp. 863-868, 1972. Huntsville, May 2014. 19. G. Liu and M. Liu, Smoothed Particle Hydrodynamics: A Meshfree 7. G. Romanelli, A. Mignone and A. Cervone, "Pulsed fusion space Particle Method, 2003. propulsion: Computational Magneto-Hydro Dynamics of a multi-coil 20. J. J. MacFarlane, I. E. Golovkin and P. R. Woodruff, "HELIOS-CR – A parabolic reaction chamber," Acta Astronautica, vol. 139, no. October, 1-D radiation-magnetohydrodynamics code with inline atomic kinetics pp. 528-544, 2017. modeling," Journal of Quantitative Spectroscopy and Radiative Transfer, 8. J. Cassibry, R. Cortez, C. Cody, S. Thompson and L. Jackson, "Three vol. 99, no. 1-3, pp. 381-397, 2006. Dimensional Modeling of Pulsed Fusion for Propulsion and Terrestrial 21. J. Monaghan, "Smoothed particle hydrodynamics," Reports on Progress Power Using Smooth Particle Fluid with Maxwell Equation Solver in Physics, vol. 68, pp. 1703-1759, 2005. (SPFMaX)," in 53rd AIAA/SAE/ASEE Joint Propulsion Conference, Atlanta, GA, 2017. 22. I. Lindemuth and R. Siemon, "The fundamental parameter space of controlled thermonuclear fusion," American Journal of Physics, vol. 77, 9. J. J. Monaghan, "Smoothed Particle Hydrodynamics," Reports on no. 5, pp. 407-416, 2009. Progress in Physics, vol. 68, pp. 1703-1759, 2005. 23. I. R. Lindemuth and R. C. Kirkpatrick, Nuclear Fusion, vol. 23, p. 263, 10. J. P. Morris, J. F. Patrick and Y. Zhu, "Modeling Low Reynolds Number 1983. Incompressible Flows Using SPH," Journal of Computational Physics, vol. 136, pp. 214-226, 1997. 24. A. M. Buyki and et al, "Investigations of Thermonuclear Magnetized Plasma Generation in the Magnetic Implosion System MAGO," 11. J. H. VanSant, "Conduction Heat Transfer Solutions," UCRL-52863-Rev. All-Russion Institute of Experimental Physics, Arzamas-16, Nizhny 1, Lawrence Livermore National Laboratory, CA, 1983. Novgorod Region, Russia. 12. R. Fatehi and M. T. Manzari, "Error estimation in smoothed particle 25. J. T. Cassibry, M. A. Rodriguez and K. J. Schillo, "Test Suite for hydrodynamics and a new scheme for second derivatives," Computers Hydrodynamic Modeling for Plasma Driven Magneto-Inertial Fusion," and Mathematics with Applications, vol. 61, no. 2, pp. 482-498, 2011. in 52nd Annual AIAA/SAE/ASEE Joint Propulsion Conference, Salt Lake City, UT, July 2016.

Received 15 May 2018 Approved 25 September 2018

JBIS Vol 71 No.4 April 2018 125 JBIS VOLUME 71 2018 PAGES 126–129

FISSION FRAGMENT ROCKET: Fuel Production and Structural Considerations PAULI ERIK LAINE, Departments of Physics & Computer Science and Information Systems, University of Jyväskylä, 40014, Finland email [email protected]

Fission reaction provides about 200 MeV of energy per nuclei. This is high compared to fusion, e.g. 1.44 MeV in 1H + 1H reaction. Uranium fission releases 81% of its energy in form of kinetic energy. Typically, this kinetic energy of the fission fragments is dissipated by collisions with other atoms to produce heat. However, the fission fragment escape probability increases as fission fuel particle size decreases, drastically when the size approaches about 10 microns. This also requires that fissile material should be ultra-thin or in low density state. This limits usable material to highly fissionable nuclear fuels, such as americium (Am) or curium (Cm), which are very expensive to produce. Chapline [1] proposed an idea of using escaping fission fragments as rocket propulsion. Since then, the idea of fission fragment rocket has been developed further (e.g. [2]). However, some fundamental constraints remain to be solved if such rocket engine could be used in an ideal interstellar mission. These questions include problems such as how to produce enough Am or Cm fuel, how to solve certain structural and thermal design issues and how to use nuclear rockets in a safe way.

Keywords: Fission Fragments, Propulsion, Am-242m, Thermal Limit, Fuel Supply

1 INTRODUCTION

Fission fragment rocket engine (FFRE) is a proposed and TABLE 1 Critical masses of some fissile promising propulsion technology due to its high exhaust ve- isotopes, from [2] locity (3-5% c) and Isp (~106 sec) [3]. Fission reaction splits atoms into two fragments that dissipate their kinetic energy Fissile Cross-section Mass* Isotope (kg) by colliding into other atoms. For example, uranium fission σpv releases 81% of its energy in form of kinetic energy that then Am242m 4.27 x 105 0.5 heats the fissile material. Theoretically, FFRE can harness these fragments’ kinetic energy directly to produce thrust. In order Cf251 1.65 x 105 0.9 for fission fragments to escape, fissile material should be ul- 245 4 tra-thin or in low density state. This limits usable material to Cm 9.34 x 10 1.1 some highly fissionable isotopes. Table 1 shows critical masses Pu239 4.04 x 104 5.6 of some fissile isotopes that uniformly fills a cylindrical core with a diameter of 1 m and a length of 5 m inside BeO or heavy U235 2.31 x 104 11.0 water neutron moderator.

The most probable mechanisms for fission fragment escape from the fissile material are recoil, diffusion, and chemical at- tack [4]. For the FFRE, the recoil is the most important escape mechanism. The relation for computing fractional release by recoil from spherical fuel particles is: probability Where fR is the fraction of fission fragments produced that are released by recoil, R recoil range (6 microns in uranium), and a radius of fuel particle. The escape probability as a function of fuel particle size is shown in figure 1. thickness (microns) This paper was presented at the Tennessee Valley Interstellar Workshop 2017 Symposium, Huntsville, USA. Fig.1 Fission fragment escape probability, from [2].

126 Vol 71 No.4 April 2018 JBIS FISSION FRAGMENT ROCKET: Fuel Production and Structural Considerations

disks exhaust electron separation fissionable reactor core electrodes filaments

fissionable filaments

(MODIFIED WIKIMEDIA COMMONS’ IMAGES) decelleration and ion collection RF induction coils electrodes containment field generator moderator

Fig.2 Original fission fragment rocket Fig.3 Dusty Plasma Bed Reactor

For this, the fissile material in FFRE reactor core should be In the DPBFFR, fissile material is in the form of a magneti- in ultra-thin layer (about one free-mean-path of a fission frag- cally confined cloud of nanoparticle dust (< 100 nm). The den- ment in solid fissile fuel) of have low density in order for the sity of this dust cloud must be balanced for fragments to escape fission fragments to escape, below 10-4 g/cm3 [5]. In Table 1, and achieve criticality. Escaping fission fragments are reflected only Am242m has average density low enough in mentioned and collimated by magnetic field. core dimensions. As the core grows, the fission fragment extraction becomes more difficult and the mass of the FFRE -in In some FFRE concepts fission fragments are used to direct- creases. In order to collect and focus fission fragments, they ly heat a conventional nuclear thermal rocket propellant (hy- must be guided with magnetic fields. Few solutions for the fis- drogen) for propulsion [6]. sion fragment extraction have been proposed. 2 FUEL PRODUCTION ISSUES Chapline first proposed the idea of using escaping fission fragments as rocket propulsion [1]. In his idea, the fuel materi- Although both introduced solutions can, in principle, operate al is in the form of thin (2 mm) carbon fibers coated with fissile with more abundant isotopes, such as plutonium, increased material on rotating disks (Fig. 2). critical mass will affect the structural design. Contrarily, isotopes such as americium and californium are difficult and -ex Only a small portion of fuel is inside the reactor core at the pensive to produce. For these reasons, feasible FFRE is a com- time, while other parts of the disks are radiating heat away. In promise between availability of the fuel and cost-effectiveness this solution, magnetic fissile fragment extraction is done with of the reactor system. a rather complex arrangement of external conductors. The most optimal fuel for FFRE is also the least available – In 2005, Clark and Sheldon [2] proposed an alternative de- the Am242m isotope. Nuclear reactions related to americium sign for FFRE, the Dusty Plasma Based Fission Fragment Re- generation are illustrated in Fig. 4. actor (DPBFFR) (Fig. 3). WIKIPEDIA COMMONS

Fig.4 Nuclear reactions related to americium generation and transmutation, from [7].

JBIS Vol 71 No.4 April 2018 127 PAULI ERIK LAINE WIKIPEDIA COMMONS

Fig.5 Nuclear reactions related to californium generation and transmutation.

In the nuclear wastes from light water reactors, the abun- ple, be used as a FFRE fuel. The problem with these isotopes, dance of Am242m is a low 0.4% of the overall americium as a however, is the fact that their critical masses are much larger chemical element [8]. In the region above 0.2 eV, the Am241 than americium, californium and curium. This indicates that cross-section has important resonances that are useful for more fissile material should be in the reactor to start fission; it Am242m production. This means that Am242m can be ob- could be difficult to build such FFRE’s. In fact, plutonium 239 tained by a neutron capture reaction with Am241. However, is produced as a byproduct in common nuclear reactors (Fig. 7) the problem is how to separate Am242m from Am241. The and uranium 235 is the only fissile isotope that is a primordial solution for the enrichment of Am242m could be the irradia- nuclide and can be found in nature. tion of Am241 in a fast reactor [9]. Unfortunately, there are not so many fast reactors available for this. 3 THERMAL ISSUES

Other isotopes can be used; however, while their properties Thermal issues play an important role in designing FFREs. become less favorable as FFRE fuel, they are increasingly easi- Some of the fission fragments will inevitably collide with fuel er to produce. For example, isotopes with high thermal fission material, dissipating their kinetic energy to produce heat. In cross-section are Californium 251 (Cf251) and Curium 245 both rotating fuel and DPBFFR, over 50% of the fragments are (Cm245). Only nuclear reaction charts are presented in Fig. 5 deposited in the fuel and reflector; thus, a cooling system is (Cf251) and Fig. 6 (Cm245). mandatory.

Also, plutonium or even uranium isotopes can, in princi- In Chapline’s [1] original design, heat is radiated by large disks that are outside of the reactor. In the DPBFFR design, the reactor chamber wall acts as a cooling system, partly by reflecting infrared (IR) light into the exhaust port and can be threaded with a high-temperature NaK cooling system [2]. In both cases, large external heat radiators are required. (FROM: WORLD NUCLEAR ASSOCIATION, PU-239 (FROM: WORLD NUCLEAR ASSOCIATION, probability

years in reactor

Fig.6 Nuclear reactions related to curium 245 generation and Fig.7 Plutonium production in the reactor core. transmutation.

128 Vol 71 No.4 April 2018 JBIS FISSION FRAGMENT ROCKET: Fuel Production and Structural Considerations

4 STRUCTURAL ISSUES laws and restrictions for using nuclear material, especially when they pose launch risks. Surprisingly, FFRE does not There are many different structural issues in both FFRE de- pose any particular risks after its nuclear material has been signs. These issues also depend on used fissile material. Real- launched to space. The FFRE creates a lower environmental istically, more common isotopes require larger structures. This issue than a nuclear thermal reactor or a space nuclear reactor is especially true in Chapline’s [1] system: rotating disks should needed for fusion propulsion [8]. In contrast, with a conven- be massive to enable fission and reflect heat. They should also tional nuclear reactor, the FFRE fission fragments are contin- rotate considerably fast, which causes further structural prob- uously expelled from the core at high velocity and leave the lems. In both designs, large magnets are required to produce vicinity of the reactor. magnetic field strong enough to extract and focus fragments for thrust. There are also numerous issues regarding which ma- 6 CONCLUSIONS terials could stand the heat and other stress. FFRE is a proposed and promising propulsion technology due 4.1. Refueling to its high exhaust velocity. Two basic FFRE designs have been proposed: a rotating disk reactor and a DPBFFR. The leading One especially important structural issue is how the new fis- question regarding a feasible FFRE is the availability of optimal sile material could be supplied into the reactor to sustain the fuel, e.g. Am242m, which will hopefully be answered sooner critical mass. How much specific impulse can be acquired if rather than later. Although there have been promising discov- original fuel is used? Probably not enough for any deep space eries, there are still basic, physics analyses (as opposed to new mission. Should there be intervals for refueling and how could physics yet to be discovered) that remain before the viability this be accomplished? Obviously, the DPBFFR should be easier of FFRE. to refuel. Acknowledgements 5 SAFETY ISSUES The author thanks Robert Werka at Marshall Space Flight Using nuclear material raises safety issues. There are many Center (MSFC) for his discussions.

REFERENCES 1. G. Chapline, “Fission Fragment Rocket Concept”, Nucl. Instr. and 6. C. Rubbia, “Fission fragments heating for space propulsion,” CERN SL- Meth., A271, pp 207-208, 1988. Note2000-036 EET, 2000 2. R.A. Clark and R.B. Sheldon, “Dusty Plasma Based Fragment Nuclear 7. M. Osaka, S. Koyama, S. Maeda and T. Mitsugashira, “An Experimental Reactor”, in 41st AIAA/ASME/SAE/ASEE Joint Propulsion Conference & Investigation of Accumulation and Transmutation Behavior of Exhibit, Meeting Papers, 2005. Americium in the MOX Fuel irritated in a fast reactor”, Annals of 3. R. Werka, R. Clark, R. Sheldon and T. Percy, “Final Report: Concept Nuclear Energy, 32, pp.635-650, 2005. Assessment of a Fission Fragment Rocket Engine (FFRE) Propelled 8. P. Bennetti, A. Cesana, L. Cinotti, G.L. Raselli and M. Terrani, Spacecraft”, FY11 NIAC Phase 1 Study, 2011. “Americium 242m and its Potential Use in Space Applications”, J. 4. J.W. Prados and J.L. Scott, “Models for Fission-Gas Release from Coated Physics: Conference Series, 41, pp.161-168, 2006. Fuel Particles”, ORNL-3421, Oak Ridge National Laboratory, 1963. 9. Y. Ronen and E. Shwageraus, “Ultra-thin Am242m fuel elements in 5. G. Chapline and Y. Matsuda, “Energy Production Using Fission nuclear reactors”, Nuclear Instruments and Methods in Physics Research, Fragment Rockets”, in Proceedings of the ICENES ’91 Conference on 455, pp.442-451, 2000. Emerging Nuclear Energy Systems, Monterey, CA, 1991.

Received 14 May 2018 Approved 25 September 2018

JBIS Vol 71 No.4 April 2018 129 JBIS VOLUME 71 2018 PAGES 130–132

FLYING ON A RAINBOW A Solar-Driven Diffractive Sailcraft GROVER A. SWARTZLANDER, JR. Chester F. Carlson Center for Imaging Science, Rochester Institute of Technology, Rochester, NY 14623 USA email [email protected]

Radiation pressure afforded by natural broadband sunlight upon a transmissive diffractive sail is theoretically and numerically investigated. A grating period of one micrometer is found to convert 83% of the solar black body spectrum into sailcraft momentum. Non-optimized orbit-raising trajectories for diffractive and reflective sails are compared. Potential advantages of diffractive sails are also described.

Keywords: Solar sailing, Sailcraft momentum, Diffraction and reflection

Recent and upcoming sailcraft demonstration missions are beginning to utilize the free and abundant momentum of solar photons for in-space navigation and propulsion [1-5]. An alternative to the longstanding assumption of a reflective sail [6,7] has recently been proposed [8], whereby a thin single order diffraction grating replaces the metal-coated membrane. The net force on a single order diffraction grating owing to the broadband solar spectrum is reported here for a transmissive grating that is suitable for orbit-raising or inclination cranking. Combining the grating equation and the solar black body spectral exitance, pressures that are comparable to reflective sails are derived; yet, with different forcing laws (e.g. the force may be tangential to a diffractive film). Solar-driven sails have been proposed for missions ranging from the near-Earth to the Fig.1 Light incident upon a diffraction grating of period Λ, with interstellar realms [9]. A simple proof-of-concept example is the sunline and surface normal unit vectors subtending the reported here whereby the sail makes a synchronous transfer angle θi. Incident, diffracted, grating wave vectors: . orbit between Earth and Mars. Such transfers may occur with- Tangential unit vector, . Azimuth unit vector, . Diffraction out changing the relative attitude of the sail with respect to the angle, . Radiation pressure force . sunline if the force in the orbital direction is greater than the force in the sunline direction [8]. While both diffractive and defined in Fig. 1. The diffraction angle θm may be expressed in reflective sails may satisfy this condition, the former brings terms of sine and cosine functions: additional advantages such as electro-optic beam steering [10] and the reuse of transmitted light (photon recycling), whereas (1) the latter provides only electro-optic diffusivity [1]. where the plus (minus) sign corresponds to a reflected (trans- A beam of light with wavelength λ is diffracted from a grat- mitted) grating order, and the expression on the left is the grating having a period Λ according to a phase matching condition ing equation which is readily derived from the above boundary at the grating boundary, as dictated by Maxwell’s equations (e.g., condition. Faraday’s Law requires continuity of the tangential electric field at the dielectric interface). Assuming the grating momentum For a given wavelength the radiation pressure force owing to vector is tangential to the surface with magnitude , a grating may be expressed [11]: the boundary condition may be expressed where and are (2) the incident and mth order diffracted wave vectors respectively, and are the respective parallel and normal unit vectors of the grating surface, and the angles θi and θm are where , is the fraction of incident beam power that is diffracted into the mth order, I is the irradiance, A is the area of the grating and c is This paper was presented at the Tennessee Valley Interstellar the speed of light. For orbital dynamical modeling it is conven- Workshop 2017 Symposium, Huntsville, USA. ient to represent the force with respect to the optical axis (i.e.

130 Vol 71 No.4 April 2018 JBIS FLYING ON A RAINBOW A Solar-Driven Diffractive Sailcraft the sunline) coordinate system, where and . Furthermore, if the grating is exposed to a beam having a spectral irradiance distribution , then the net force may be determined by integration:

(3a)

(3b) where is the wavelength-dependent mth order diffraction efficiency. IfM(λ) represents the solar black body spectral ex- Fig.2 Black body solar spectral irradiance at T=5778K (solid line). itance, then where Transverse (circles) and longitudinal (diamonds) solar radiation is the solar radius, r is the distance between the sun, the sail- pressure spectral density distribution for an ideal single order craft and , where h grating of period Λ=1 μm and incidence angle θi=20°. and kB are respectively the Planck and Boltzmann constants and T=5778K is the effective temperature of the sun. The solar 11 spectral irradiance at a distance of r = 1[AU](1.496 x 10 [m]) is where RE = 1 [AU], G is the universal gravitational constant, shown in Fig. 2. Although the sun is not a perfect black body ra- while and M are the respective sailcraft and solar masses. The diator, integration of B(λ) across all wavelengths at 1 AU agrees equation of motion may be expressed in normalized form: with the measured value of the “solar constant” . (5) The allowed diffraction angles, dictated by Eq. (1), provide lower and upper wavelength bounds: and where is the sailcraft areal den- , respectively. Examples of the spec- sity, is the so-called lightness number, and tral diffraction range are evident in Fig. 2 for a grating pe- is a characteristic areal density. riod of , with a diffraction cut-off wavelength at Note that as the areal density σ decreases, the lightness increases. . This period is an optimized value that makes use of roughly 83% of the available solar power. In comparison, Assuming circular co-planar planetary orbits, a synchro- reflective sails may use a similar fraction after accounting for nous transfer is one that matches the boundary conditions be- absorption and re-radiation, as well as the trade-off between tween planet A and planet B: the metallic layer thickness and its mass.

In practice, modern gratings may be designed with me- where ta-material engineering approaches to optimize the diffraction efficiency across a band of wavelengths. Below, an ideal grating and is the transit time determined from numerical integra- having 100% diffraction efficiency into the m=1 order is as- tion of Eq. (5). The sail is jettisoned or stowed at to maintain sumed for all wavelengths up to λmax. The solar radiation pres- the desired final orbit for . Synchronous or quasi-syn- sure force on a grating was numerically computed for the solar chronous transfer orbits may be achieved from Earth to Mars black body spectrum. Optimizing for both a large magnitude if and 0.0635[8]. To spiral outward, the grating of transverse force and a small value of longitudinal force, Λ=1 may be designed to have a dominant diffraction orderm =-1, [μm] and θi = 20° was determined, along with the correspond- in which case the attitude is set to θi = –20° with respect to ing radiation pressure spectral density distributions shown in the sunline, resulting in a transverse (radial) force efficiency Fig. 2. Integrating across wavelengths provides the values of the value of =+0.5 ( =+0.22). The transfer requires a sailcraft net force components per unit area: areal density of roughly 12.1 [g/m2] (i.e. =0.127). Using the and at r = 1 AU. The corresponding fourth-order Runge-Kutta numerical integration technique force efficiencies may be expressed and [12], a value of =1.439 years (525.6 days) is found for an . This is a desired condition for changing Earth to Mars transfer. The transfer time was similarly cal- the orbit or inclination, i.e. . culated to be four days longer for a reflective sail also having =+0.5. In comparison, the force efficiencies of an ideal non-absorbing, perfectly reflecting sail are given by and In principle, any mission trajectory must allow adjustments . In this case the incident angle must of the force components to account for disturbances or to op- exceed 45° to achieve the criterion . This does not timize the transfer time. Such a detailed study is beyond the make efficient use of the sail because the available driving pow- scope of this report. Like a reflective sail, the force on a diffrac- er on the sail falls as . Nevertheless, the reflective sail tive sail may be varied by changing the attitude of the sail with achieves a transverse efficiency when θi = 56.9° with respect to the sunline. In this case, the forcing law of a sun-lit a corresponding longitudinal efficiency of 0.33. The latter -val grating having a period Λ=1 μm is shown in Fig. 3 overleaf. The ue is larger and therefore less desirable than the corresponding first diffraction order (m=1) provides the greatest magnitude value (0.22) for the diffractive sail discussed above. of transverse force. What is more, by varying the attitude, say between -20° and 20°, the transverse force may be changed by The net force on the sailcraft from solar gravity and radiation roughly 50%. Note that negative diffraction orders produce the pressure at any distance from the sun, r, may be expressed [9]: same lines, just with and .

(4) One of the potential advantages of diffractive sails is the op-

JBIS Vol 71 No.4 April 2018 131 GROVER A. SWARTZLANDER, JR.

Fig.3 Net transverse (solid lines) and longitudinal (dashed lines) Fig.4 Net transverse (solid lines) and longitudinal (dashed lines) force efficiencies of a solar-driven diffractive sail having a grating force efficiencies of a solar-driven diffractive sail having a variable period Λ=1 μm for diffraction orders m=1,2,3. grating period for diffraction orders m=1 and -1. portunity to change the diffractive properties by electro-optic may therefore afford a route to interstellar space [14,15]. means [10]. This may allow a forcing law based on photonic rather than cumbersome mechanical control systems. Two dif- In summary, the premise that sun-driven diffractive sails ferent approaches to this solution may be described with the may replace reflective sails to achieve mission objectives, such aid of Fig. 4, which plots the force efficiencies for the m=1 and as orbit raising, has been theoretically validated. The same ar- -1 orders of a solar driven sail having an attitude of 20°. By gument may be made for orbit lowering (e.g. an Earth to Venus actively varying the grating period, components of force may trajectory) and inclination cranking. A grating period of Λ ≈ 1 be directly controlled. In particular, the transverse force effi- μm was found to achieve a large transverse efficiency across the ciency is found to vary linearly when the grating period varies solar spectrum. The principle of non-mechanical navigation between Λ = 1 and 2 μm (see straight, dashed line in Fig. 4). methods were proposed whereby either the grating period or Another proposed force control scheme involves switching the grating order is varied, e.g. via electro-optic means. Photonic grating order from +1 to the -1. Changing the sign of the order and meta-materials research is needed to develop the required changes the direction of the transverse force, thereby allowing diffractive films that provide a high efficiency single diffraction navigation of the sailcraft. order across the visible and near-infrared region of the solar spectrum. The film must also be able to withstand the harsh Another advantage of a diffractive sailcraft is uncovered by ultraviolet environment of outer space. the fact that the sail transmits photons that may be re-used to enhance the radiation pressure force or to generate photovoltaic Acknowledgements energy for the communication and control bus. What is more, This research was partially supported by the NASA Innovative absorption within the diffractive sail may be negligible, thereby Advanced Concepts Program (NIAC) (80NSSC18K0867). The obviating heat stress and re-radiation problems of metallic films author is grateful to Les Johnson, NASA Marshall Space Flight [9, 13]. Unfurling a diffractive sail in close proximity to the sun Center, for valued discussions about solar sailing.

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Received 15 May 2018 Approved 25 September 2018

132 Vol 71 No.4 April 2018 JBIS JBIS VOLUME 71 2018 PAGES 133–139

EVALUATION OF THE HAZARD OF DUST IMPACTS on Interstellar Spacecraft RICHARD A. LONDON1 and JAMES T. EARLY2 1Lawrence Livermore National Laboratory, Livermore, CA, 94550, USA. 27056 Currituck Rd, Kitty Hawk, NC 27949, USA. email [email protected] / [email protected]

Spacecraft traveling to nearby stars will experience energetic collisions with interstellar dust grains. For thin structures, such as light sails, grains pass through with little damage. For thick structures, all the grain energy will be deposited, possibly leading to significant damage. This paper presents a quantitative assessment of damage caused by grain-vehicle collisions and designs to reduce damage. Computer simulations of grain-vehicle collisions were performed with a radiation/ hydrodynamic code, that models deposition of grain energy in the material, energy transport by conduction and radiation, expansion of the heated “volume melting” and other damage processes to the surrounding material. It is found that the grain energy is deposited in a long, thin cylindrical volume, creating temperatures larger than 107 K. The spreading of this energy can potentially lead to a large damaged volume. Radiation and evaporation can remove energy, thereby limiting the damage. The reduction of damage by the choice of shielding material on the leading surface and by placement of a thin shield in front of the vehicle to atomize and disperse the grains before they hit the main body of the craft is described.

Keywords: Interstellar travel, Dust, Grains, Damage

1 INTRODUCTION puter simulations to the study of collisions of dust grains with interstellar spacecraft. Space missions to nearby stars will require travel at quasi-relativistic speeds in order to reach the destinations and return 2 PROPERTIES OF NEARBY INTERSTELLAR MATTER information in a reasonable amount of time. For example, a speed of 0.2 of the speed-of-light is needed for a 25-year mis- Information on interstellar matter is derived from two sources: sion to the nearest star system, Alpha Centauri (α Cen), at a astronomical oservations of nearby stars and direct measure- distance of 4.4 ly. At such high speeds, collisions with inter- ments of dust grains by solar system spacecraft. stellar matter could cause considerable damage to a spacecraft. Previous work has examined collisions of interstellar dust Astronomical observations consist primarily of the effects on grains with light sails used for laser-driven acceleration [1, 2]. starlight, such as visible extinction, reddening, polarization and That analysis estimated that the damage area by a single grain infrared absorption, and emission from nebulae and other ex- was not much larger than the cross-sectional area of a grain it- tended objects, as reviewed by Draine [7]. Grains are thought to self. Given the size and number of grains that would be expect- be of two types–carbonaceous, with densities of approximately ed to be encountered during the vehicle acceleration phase, 2 g/cm3 and composed mainly of C, and silicate, with densities 3 the total damage to a light sail would not be large. A leading of 3.5 g/cm and a typical composition of MgSiO4Fe. The main thin foil combined with an electric field on the main body of focus is on silicate grains, since their mass in the interstellar a craft was proposed to atomize and ionize grains and deflect medium (ISM) is more than twice that in carbonaceous grains. the ions, thereby protecting large spacecraft [2]. Recently, the A typical size distribution of silicates [8] is illustrated in Fig. 1 concept of near-earth rapid laser acceleration of extremely overleaf. Observations suggest that large fractions of the heav- small (cm-scale) interstellar probes has been proposed [3,4]. ier elements (Mg, Si…Fe) are depleted from the gas phase and Estimates of the damaging effects of collisions with interstellar bound up in dust grains. The mass density of silicate grains can atoms and dust grains on such small probes have been made be estimated from the hydrogen density in the local ISM and by Hoang, et al. [5]. The most important damage mechanism assuming solar abundances for the elements. Slavin and Frisch is thought to be collisions with interstellar dust. Collisions can [9] give a typical value for the hydrogen density in the local ISM also cause rotation and deflection of a probe by direct momen- of 0.25 cm-3. Specific spectroscopic observations along the line- tum transfer and by charging and interaction with the inter- of-sight to α Cen [10] give values for the column densities of gas stellar magnetic field [6]. In this paper, the study of damage phase deuterium (D), Mg and Fe. Taken together with a ratio of effects is extended by applying radiation/hydrodynamic com- D/H in the solar neighborhood (the “local bubble”) of 1.56 x10- 5 [11], the D measurements indicate a column density of H of 4x1017 cm-2 and an average density of 0.09 cm-3. They also meas- This paper was presented at the Tennessee Valley Interstellar ured a large depletion of Mg and Fe from the gas phase. Their Workshop 2017 Symposium, Huntsville, USA. H density is somewhat smaller than suggested in ref. [9]. This

JBIS Vol 71 No.4 April 2018 133 RICHARD A. LONDON & JAMES T. EARLY

less subject to damage than most insulating materials, an effect ascribed primarily to the higher degree of exciton self-trapping in the latter materials. Hoang, et al. estimate that approximately 30% of a quartz surface would be damaged in traversing a hydrogen column density of ~ 4x1017cm-3 at a speed of 0.2 c, while a graphite surface would not be significantly damaged.

4 EFFECTS OF GRAIN-SPACECRAFT COLLISIONS

4.1 Summary of prior work

Hoang, et al. [5] also considered the effects of dust grain impacts. They estimated the volume of material that could be evaporated and melted by equating the dust grain energy to the threshold internal energy required for those processes and factoring in the overlap of damaged regions by multiple impacts. They estimated that more than 0.5 mm of a quartz surface could be evaporated, and more than 3 mm could be melted in traveling to α Cen. These estimates did not consider Fig.1 Column density of grains larger than a given radius between several details of the dust grain-spacecraft collisions, such as the earth and α-Cen. The black lines are based on an interstellar radiative energy loss, mass loss at higher specific energy than model of ref. [8], Eq. (5) with parameters from their Table 1 the evaporation threshold, energy transfer to regions outside -5 with RV=3.1 and bc= 6x10 . The 2 curves are for limiting atomic the melt volume, motion of the melted material and mechani- hydrogen densities, as discussed in the text. The red bars are cal damage mechanisms, such as crushing and spallation. from Ulysses measurements [12] where the heights indicate the uncertainty range. 4.2. Energetics of dust grain collisions

In the reference frame of the spacecraft, each dust grain carries contrast may represent differences in the sampled lines-of-sight. a large energy: The range in densities is treated as an uncertainty. Assuming that most of the silicate elements are condensed into grains and that their abundances are solar, these values of H density yield (1) estimated mass column densities of dust to α Cen between 0.8 and 2.2x10-8 g/cm2.

Measurements of interstellar grains that penetrate the he- where ρ and r are the grain density and radius, respectively, liosphere have been made with interplanetary spacecraft. A and β is the spacecraft velocity relative to the speed of light. new compilation and analysis of data from the Ulysses space- This formula ignores the relativistic correction, which is less craft has been given by Kruger et al. [12]. Interstellar grains than 3% for β < 0.2. This energy will be imparted to the outer were separated from those of interplanetary origin with sever- layers (0.1 to 1 mm) of the leading surface of the spacecraft. al selection criteria and found to be 15% of the total. A mass A primary process leading to damage and/or mass loss of density of dust associated with the local interstellar medium spacecraft material is melting. To melt any particular material, of 2x10-27 g/cm3 was determined. This implies a dust column enough energy must be imparted to raise the temperature to density to α Cen of 0.9x10-8 g/cm2, within the range inferred the melt value and then to cause the phase change from sol- from astronomical observations. As evident in Fig. 1, there id to liquid. The energy density required to melt typical solid are notable difference between the grain distributions inferred materials that might be used on the leading edge of the craft from astronomical observations and those measured by Ul- is of order 1 eV/atom. For example, the melt energy is 2.2, 0.9 ysses. The difference at small radius (<0.2 µm) is likely caused and 1.6 eV/atom for diamond, silicon and tungsten, respec- by deflection of charged grains by the solar magnetic field [ref. tively [14]. Therefore, a grain with parameters given in Eq. (1) 12 and references therein]. The density of large grains is rather could melt 1.5 mg of diamond, 8.6 mg of silicon or 32 mg of uncertain due to small number of detections at these sizes. The tungsten. Assuming a mass column density of dust to α-Cen Ulysses measurements, as well as prior measurements with the at the low end of the estimates given above, the total ener- Galileo spacecraft, indicate more grains at large sizes than the gy per unit area imparted to a spacecraft traveling to α-Cen astronomical observations (Fig. 1). Given the estimated den- is 14 kJ/cm2. If all of this energy goes into melting, it would sity of larger grains from the spacecraft measurements, such melt approximately 0.8, 4.7, or 17 g/cm2 of diamond, silicon collisions with interstellar craft must be considered. or tungsten, respectively. Since spacecraft masses are desired to be only several grams, such melting could be catastrophic. 3 EFFECTS OF GAS-SPACECRAFT COLLISIONS This situation is possibly remedied by removal of energy by radiation and/or high velocity blowoff. Ionization of the blow- Hoang, et al. [5] recently analyzed the effects of collisions of off, as observed in dust particle collisions with interplanetary interstellar gas-phase atoms with spacecraft material. Under spacecraft [15] and in laboratory experiments [16], will en- certain conditions, such collisions form nm-wide by mm-long hance the energy removal. Dispersion of the deposited energy damage tracks. They estimate that the most damaging collisions into a large volume, such that most of the volume receives a are with O and Fe atoms, due to the combination of a large lower specific energy than required to melt and refreezing of effect per atom and relatively large abundances. As discussed melted material before it has time to leave the spacecraft may by Itoh, et al. [13], metals and crystalline semiconductors are also reduce the amount of spacecraft material that is melted.

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Fig.2 Schematic diagram showing the time sequence of the collision Fig.3 Energy loss rate versus temperature. Results are given for 4 of a dust grain with the material at the front surface of a spacecraft. combinations of grain and spacecraft materials, as labeled.

4.2. Collision length and time scales rials (Si and W). These calculations assume steady-state local thermodynamic equilibrium (LTE) ionization of the target Fig. 2 illustrates the characteristic progression of the collision material. Fig. 3 illustrates a decreasing energy loss rate at high of a single dust grain with a spacecraft. Moving atβ = 0.2, a temperature for all cases. The rate is highly dependent on the micron-scale grain will pass into the craft in about 20 fs. Due to atomic number of both the grain and target elements. The en- energetic collisions, the atoms of the grain will become highly ergy loss rate at low temperature ranges from 0.3 GeV/mm for ionized. The atomized and ionized grain material penetrates O on Si to 15 GeV/mm for Fe on W. The reduction in energy further into the spacecraft, heating each layer of the spacecraft loss rate at high temperature leads to a larger penetration depth on a similar timescale. The grain matter will be stopped in 0.05 in the calculations, compared to those of ref. [5], where cold to 1 mm, within a time of 1 to 20 ps. During a much longer pe- energy loss rates were. riod of time, the deposited energy will spread via a high-pressure hydrodynamic wave (initially a strong shock wave), x-ray 4.4. Numerical Simulations of grain-spacecraft collisions and UV radiation as well as heat conduction. During this period, some of the energy will be removed from the craft by radi- 4.4.1. Techniques ation and high velocity mass loss. It is estimated that the final Numerical simulations were performed to quantify the energy state of the material will be reached in 1-10 µs, when vaporized deposition and resultant effects, using the high energy-densi- and melted material are removed and the remaining material ty radiation-hydrodynamic code HYDRA [19]. HYDRA has cools below the melt temperature. been developed to model inertial confinement fusion experiments as well as extensively benchmarked against experi- 4.3. Energy deposition mental data [20, 21]. Physical variables, such as temperature, density and ionization are defined on a spatial mesh and ad- The interaction of the grain with spacecraft material occurs vanced in finite time steps. The heavy-ion beam target irra- by fast atomization and stripping, then stopping of the result- diation package was used to model the dust grain-spacecraft ing ions. The primary stopping mechanism is collisions with interaction. This model calculates a self-consistent ionization target electrons, both bound and free. The following formula state of the grain atoms according to the Betz theory [17, 18]. is used for the energy loss rate (i.e. stopping power) of ions The ions propagate on straight lines, transferring energy to the [17,18]: target material according to the energy loss rate given by Eq. (2). The temperature, ionization state and pressure of the target matter are calculated self-consistently with time-depend- (2) ent non-LTE rate equations. Hydrodynamic motion – particularly expansion of the heated material and compression of the material outside the heated region by pressure driven shock where Ei is the energy of the grain ion, z is the distance along waves – is simulated by finite difference equations and treated the ion path within the spacecraft, e and m e are the electron with the Lagrangian method for one-dimensional problems charge and mass, c is the speed of light, Zg is the charge of the and the arbitrary Lagrangian-Eulerian (ALE) method for grain ion, ns is the number density of spacecraft atoms, Zb and two-dimensional problems. Radiation transfer, both into the Zf are the number of bound and free electrons per spacecraft adjacent material and into space from the surface, is simulat- atom and lnΛb and lnΛf are the bound and free electron Cou- ed with multi-group diffusion equations, including opacities lomb logarithms. The function G(x) = erf(x) – x erf’(x), where and emissivities derived from the non-local-thermodynamic erf is the error function. equilibrium (non-LTE) atomics physics. Heat transport by conduction (mainly electron conduction) is calculated with a Fig. 3 shows the energy loss rate as a function of the heated diffusion equation, using temperature and density dependent spacecraft material forβ = 0.2 ions of two characteristic grain plasma-based conductivities. The radiation and heat transport elements (O and Fe) and two characteristic spacecraft mate- equations are solved by the finite difference method.

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TABLE 1 Table 1. 1-D Simulation cases

Case Zs (mm) rc (mm) S = Zs/rc

A (O on Si) 1.15 0.31 3.7 B (Fe on Si) 0.50 0.37 1.4 C (O on W) 0.17 0.32 0.53 D (Fe on W) 0.07 0.33 0.21

4.4.2. 1-D simulations Since the region in which the grain material is stopped is typically very thin and long, order 10-3 mm by 1 mm, and the difference in timescales for grain stopping and for the damage processes is many orders of magnitude, a lot can be learned about the interaction with 1-D simulations. Two types of 1-D simulations were performed: 1) longitudinal simulations to analyze the details of the stopping of the grain material and the resultant initial energy density and temperature of the craft Fig.5 Temperature versus depth just after grain stopping. material and 2) radial simulations to study the mass and energy Results are given for the 4 combinations of grain and spacecraft transport perpendicular to the impact direction. Four simula- materials, as labeled. The grains are 0.3 μm in radius and are tions of each type have been performed, with permutations of travelling at β=0.2. O and Fe grain composition and Si and W spacecraft composition. These cases, labeled A-D, are listed in Table 1. The two in the energy loss rate of the initial ions as they slow down, at grain elements represent the lightest and heaviest of the atoms low temperature. As the material heats up, the stopping distance contained in silicate grains and also the major mass compo- grows, and the deeper regions of the spacecraft get heated. All nents of these grains. A grain density of 3.5 g/cc is assumed for of the grain energy is deposited within 1.2 mm, indicated by the all four cases, typical of the expected silicate density. In doing drop in temperature to nearly zero at that point. so, the mass (and thereby kinetic) energy of the grain is kept constant while varying its composition. Future simulations Fig. 5 shows the temperature profiles for all four cases at the will be done with more realistic silicate compositions. The two end of the simulations, when the grain atoms have completely spacecraft materials represent typical light and heavy elements stopped. This is the maximum temperature reached in the sim- that might be used. ulation for each case. All the grain kinetic energy goes into internal energy of the spacecraft material. Of this energy, approx- 4.4.2.1. 1-D Longitudinal Simulations imately 77% goes into heat in Si and 44% in W, with the rest For longitudinal simulations, the spatial domain is a homoge- going into ionization energy. Due to the increase in stopping neous cylinder with reflecting boundaries at a radius equal to power with decreasing ion velocity, the maximum temperature that of the grain. The stopping of the grain material and the as- occurs inside the target. Extremely high temperatures–ranging sociated heating of the spacecraft material versus depth is stud- from 15 keV for O/Si to 60 keV for Fe/Au are reached immedi- ied. The team typically uses 200 spatial zones in the longitudinal ately after the collision. Complete stopping occurs in distances (“z”) direction and 2,000 timesteps. Fig. 4 shows a space-time ranging from 0.07 to 1.15 mm, defined as zs and listed in Table plot of the temperature for an O grain, with rg = 0.3 µm, ρ = 3.5 1. The simulations used for the results presented in Fig. 4, 5 g/cm3 and β = 0.2 hitting a Si target. Time is measured from the include time dependent non-LTE ionization of the spacecraft moment that the leading edge of the grain passes each point in material. This is important since the ionization lags behind space. The temperature increases with time as the grain atoms LTE values due to the fact that the duration of the heating pulse continue to deposit energy into the spacecraft material for 10 fs. (10 fs in these cases) is faster than the collisional ionization The spike in temperature near 0.5 mm results from a steep rise time (of order 100 fs). The energy loss rates and temperatures are less with non-LTE ionization.

4.4.2.2. 1-D Radial Simulations Radial simulations have been performed to study the energy transport after the grain stops. Only a single zone in the longitudinal direction is included, with length equal to the stopping distance of the ions (see zs in Table 1). The energy deposition and resulting temperature are thus averages over the expected longitudinal variations (Fig. 5). There are 200 zones in radius included, with small widths at small radii in order to resolve the grain stopping region as well as larger widths at large radii to cover the late-time effects.

Fig. 6 shows the central temperature versus time for the four cases. In all cases, the maximum central temperature is lower than the average temperature reached in the longitudinal simulations due to spreading of the deposited energy by heat con- Fig.4 Temperature versus depth and time for O grain atoms duction. The central temperatures drop from values of 10-50 colliding with Si spacecraft material at β =0.2. keV to values near 1 eV in the µs time frame of the simulations.

136 Vol 71 No.4 April 2018 JBIS EVALUATION OF THE HAZARD OF DUST IMPACTS on Interstellar Spacecraft

Fig.6 Central temperature versus time just after grain stopping for four cases of grain and spacecraft material. The grains have radius 0.3 μm and velocity β=0.2.

The cooling is due to a combination of heat conduction, hydrodynamic expansion and radiative transfer. As the central region cools, the outer regions get heated. This is illustrated by the radial dependence of the temperature at various times (Fig. 7a). At first the energy spreading is due to electron conduction. At intermediate times, a shock wave from the rapidly expanding inner region heats the outer regions. Fig. 7b shows the radial dependence of density at various times. The expansion of the central heated region and the associated shock-wave compression of the outer regions are seen. By a time of 0.1 µs, the shock wave has weakened, so that only a small amount of compression follows.

The final radius of the low-density region, 310 µm (0.31 mm) for case A shown in Fig. 7b, is identified as the crater ra- Fig.7 Temperature (a) and density (b) versus radius at decade dius, rc, and is listed in Table 1. The crater aspect ratio, i.e. the intervals in time ranging from 1 fs to 1 μs. Several of the curves ratio of length to width, S=zs/rc, is also listed. are labelled with time. The minimum radius increases with each time step, as the inner Lagrangian zone moves outward. The The amount of material that is melted in the simulations is grains have 0.3 μm radius and β=0.2 velocity. estimated by considering the space and time dependence of the energy density, shown in Fig. 8 for case A. The white contour time. The mass of melted material versus time is shown in Fig. shows the melt energy density, i.e. 0.9 eV/atom for Si. Melted 9. The final melted mass is dependent only on the spacecraft material is identified as that for which the energy density ex- material and is about 0.06 mg for Si and 0.2 mg for W. The melt ceeds the melt energy density at any time earlier than a given masses are about 1/4 of the maximum mass that could be melt-

Fig.8 Log10 of the energy density versus time and initial radius Fig.9 Melt mass versus time from 1-D radial simulations of 4 cases, for case A. The white contour shows the melt energy density. as labeled.

JBIS Vol 71 No.4 April 2018 137 RICHARD A. LONDON & JAMES T. EARLY ed for the specified grain parameters (r=0.3 µm, ρ=3.5 g/cc, β made to begin such simulations. Rather than simulating the = 0.2). The rest of the energy is dispersed to larger radii where grain stopping and deposition with the heavy ion beam model the energy density is lower. used for the 1-D cases, the problem has been simplified by spec- ifying energy deposition in a short flat-top pulse in a region of The aspect ratio is expected to play a role in the amount radius equal to the grain radius and a depth equal to the stop- of material that is ultimately removed and/or crushed by the ping length. A further simplification has been made by limiting impact. Several competing effects come into play. One effect is the depth of numerical domain to the energy deposition depth the removal of energy from the spacecraft by radiation. This and imposing reflecting boundary conditions at that point. will reduce the energy available to damage the spacecraft material. Radiation will be enhanced by a small aspect ratio as The evolution of density for a 2-D simulation of case A is shown well as an increased atomic number of the spacecraft material. in Fig. 10. The mid-time behavior (20 to 80 ns) is similar to the Thus, W is expected to be better than Si in this respect. Anoth- 1-D radial simulations. A shock-wave is observed, moving outer mechanism favoring small aspect ratio craters is the remov- ward in radius at a velocity of about 6 µm/ns. The crater grows al of energy by high velocity blowoff. On the other hand, the to a width of about 0.1 mm. At later times (100 – 400 ns), the processes of refreezing of material will lead to less overall mass shock weakens and the crater growth slows down. At the final loss and thus favored by high aspect ratio craters. Ultimate- time of the simulation, the crater width is about 0.25 mm and ly, high accuracy 2-D simulations including material strength has developed a lip that extends slightly beyond the original will be necessary to evaluate the effects of radiation and mass surface of the spacecraft material. loss on the damage process 5 METHODS TO REDUCE DAMAGE 4.4.2.3. Parameter Variations According the Ulysses measurements of grain column densi- Several design choices can be made to reduce damage. The front ty, reproduced in Fig. 1, there might be collisions with very surface area – perpendicular to the direction of motion – should large grains. Simulations of the collision of a 10-µm grain at be minimized [3]. This will reduce the number of collisions with β=0.2 indicates that a peak temperature of approximately 70 atoms and grains and thereby reduce the required shielding keV would be created in Si spacecraft material and the grain mass. It is likely that some materials will be more resistant to ions would stop in about 1 cm. This would melt about 1.6 g of damage by grains [5]. Materials with high melt temperature and Si. Therefore, a single collision with a very large grain could high specific melt energy (per unit mass) are preferred. Materials destroy a cm-scale spacecraft. with short stopping distances can reject energy from a collision by radiation and high-velocity blow-off. There appears to be a Smaller spacecraft velocities might be advantageous to re- tradeoff between low atomic number materials, which generally duce the laser driver requirements and to reduce damage. Cal- have higher specific melt energies and high atomic number ma- culations have been performed for the collision of a 0.3-µm terials, which have shorter stopping distances and more efficient grain with Si at β=0.1. The penetration depth is reduced to 0.25 radiation. The choice of an optimal material will require further mm, compared to 1.2 mm for β=0.2 (Fig. 5). The melt mass multi-dimensional calculations, including a more careful treat- found in the radial simulation is 0.016 mg, the same fraction ment of melt dynamics and material strength effects. of the maximum melt mass as for β=0.2. These results indicate that the damage to a spacecraft scales with the grain energy, Another way to reduce damage is by placement of a thin foil which in turn scales as the square of the spacecraft velocity. ahead of the spacecraft [2]. Although it has been suggested that Therefore, damage can be significantly reduced at the expense such a foil would need to slow down the grains [5], it is neces- of longer mission time. sary only to ionize and heat the grain, allowing a much thinner foil to do the job. The electron temperature of the grain after 4.4.3. 2-D simulations such a passage from the energy loss formula [Eq. (2)] and the 2-D simulations are necessary to address questions of how foil and grain properties is estimated as: much energy is radiated away and how much is carried off in high velocity blow-off, whether material is damaged by com- (3) pressive and/or shear stress and to more accurately calculate the shape of the damage crater. Several approximations have been where nf is the number density of atoms in the foil, df is the foil thickness, and ne,g is the number density of electrons in the grain. The energy loss (dEi/dz) is now calculated for foil ions hitting grain material. The grain will explode at approx- 1/2 imately the plasma sound speed, vs=(3kZTe/M) , where Z is the average charge of the grain atoms and M is their average mass. The grain atoms will expand by a distance of vsDf /(βc), where Df is the distance between the foil and the spacecraft. This can spread the grain atoms over an area comparable to or larger than the front surface area of the spacecraft. For example, with a 20-nm Be foil placed 1 meter ahead of a spacecraft traveling at β=0.2, the grain atoms are heated to 0.3 keV and spread by about 0.3 cm. A 100-nm Be foil will spread the grain atoms by about 0.6 cm, while a 20 nm Au foil could Fig.10 2-D simulation of density vs. r and z at 8 times between 20 spread the atoms by 1.3 cm. By placing the foil at 10 m, one and 400 ns. Each panel is labeled by the time in ns. The simulation could achieve spreading of 3 to 13 cm. This spreading will is for case A, with O grain atoms hitting a Si target. The initial have two beneficial effects: many of the atoms will by spread density of the Si is 2.33 g/cc. out far enough to miss the spacecraft and the damage by the

138 Vol 71 No.4 April 2018 JBIS EVALUATION OF THE HAZARD OF DUST IMPACTS on Interstellar Spacecraft individual atoms is expected to be less than that due to a con- been performed and the results presented. densed grain. In addition, if an electrostatic potential of order 50 MeV can be imposed between the ionizing foil and a thin Damage can be reduced by optimal choice of the spacecraft foil or mesh placed just above the spacecraft, the grain ions as geometry, the material placed on the front surface and by the well as ions from initially neutral gas, could be deflected from placement of a thin foil at some distance ahead of the craft to hitting the spacecraft [2]. ionize and explode dust grains.

6 SUMMARY AND FUTURE DIRECTIONS The analysis tools described allow for the design and optimization of shields to protect interstellar spacecraft. Future This paper has described the problem of dust grain damage work is suggested to improve the 2-D calculations by incor- to spacecraft traveling at near-relativistic speeds through the porating strength in the hydrodynamics, refining numerical ISM. Given the velocity of the spacecraft and the expected resolution, and extending the domain to large depth and radi- density of dust in the local ISM, simple estimates indicate us. In addition, simulations are suggested with realistic grain that a significant mass of spacecraft material could be dam- compositions as well as with a range of grain sizes expected aged. In order to quantify the damage processes and thereby in the interstellar medium, spacecraft velocities and spacecraft refine such estimates, the team has engaged in a program of materials. Application of these simulations to a prediction of computer simulations of dust grain-spacecraft collisions. 1-D the total damage to a spacecraft in the manner discussed in ref. longitudinal simulations have been performed to quantify the [5] is suggested. This will lead to an improved assessment of the stopping of the grain and its heating of the spacecraft materi- survivability and optimized designs for the shielding layers of al with a self-consistent treatment of the energy loss rate and interstellar spacecraft. the material temperature. Radial 1-D simulations were done to study the energy transport from the initial elongated region Acknowledgements directly heated by the grain and to estimate the mass of material that is melted. These simulations indicate that the actual Many thanks to G. Basri, J. Linsky, P. Frisch and B. Wood for dis- mass of melted material is about 1/4 of the maximum value cussions of the density and composition of interstellar dust grains were all of the grain energy used for melting. Although the in the solar vicinity. This work was performed under the auspices 1-D simulations shed light on many of the physical processes of the U.S. Department of Energy by Lawrence Livermore Na- involved in the collision, 2-D simulations are needed to ac- tional Laboratory under Contract DE-AC52-07NA27344. Fund- curately determine the amount of material removed as well ing was provided by the LLNL Laboratory Directed Research as damage by the collision. A preliminary 2-D simulation has and Development Program, project 17-ERD-091.

REFERENCES 1. J.T. Early and R.A. London, “Dust Grain Damage to Interstellar Laser- Distribution and Gas-to-Dust Mass Ratio”, Astrophys. J., 812, pp. 139- Pushed Lightsail”, J. Spacecraft and Rockets, 37, pp. 526-531, (2000). 155, 2015. 2. J.T. Early and R.A. London, “Dust Grain Damage to Interstellar Vehicles 13. N. Itoh, D. M. Duffy, S. Khakshouri and A. M. Stoneham, “Making and Lightsails”, JBIS, 68, pp. 205-210, 2015. Tracks: Electronic Excitation Roles in Forming Swift Heavy Ion Tracks”, 3. P. Lubin, "A Roadmap to Interstellar Flight", JBIS, 69, pp. 40–72, 2016. J. Phys.: Condens. Matter 21, 474205 (14pp), 2009. 4. Breakthrough Starshot website, https://breakthroughinitiatives.org/3 14. R.C. Weast, Handbook of Chemistry and Physics, The Chemical Rubber (Last Accessed 3rd May 2018). Co., Cleveland, 1969. 5. T. Hoang, A. Lazarian, B. Burkhart and A. Loeb, “The Interaction of 15. A. Zaslavsky, N. Meyer-Vernet, I. Mann, A. Czechowski, K. Issautier, Relativistic Spacecrafts with the Interstellar Medium”, Astrophys. J., 837, G. Le Chat, F. Pantellini, K. Goetz, M. Maksimovic, S. D. Bale, and J. pp. 5-21, 2017. C. Kasper, “Interplanetary Dust Detection by Radio Antennas: Mass Calibration and Fluxes Measured by STEREO/WAVES”, J. Geophys. Res. 6. T. Hoang and A. Loeb, “Electromagnetic Forces on a Relativistic 117, A05102, 2012. Spacecraft in the Interstellar Medium” Astrophys. J., 848, pp. 31-38, 2017. 16. A. Collette, G. Meyer, D. Malaspina, and Z. Sternovsky, “Laboratory Investigation of Antenna Signals from Dust Impacts on Spacecraft”, J. 7. B.T. Draine, “Interstellar Dust Grains”, Ann. Rev. Astron. Astrophys., 41, Geophys. Res: Space Phys., 10.1002, pp. 5298-5305, 2015. pp. 241-289, 2003. 17. H.D. Betz, “Charge States and Charge-Changing Cross Sections of 8. J.C. Weingartner and B.T. Draine, “Dust Grain-Size Distributions Fast Heavy Ions Penetrating Through Gaseous and Solid Media”, Rev. and Extinctions in the Milky Way, Large Magellanic Cloud and Small Modern Phys., 44, pp. 465-539, 1972. Magellanic Cloud”, Astrophys. J., 548, pp. 296-309, 2001. 18. S. Atzeni and J. Meyer-ter-Vehn, The Physics of Inertial Fusion, Oxford 9. J. D. Slavin and C. Frisch, “The Boundary Conditions of the Univ. Press, 2004. Heliosphere: Photoionization Models Constrained by Interstellar and in Situ Data”, Astron. Astrophys., 491, pp. 53-68, 2008. 19. M.M. Marinak, G.D. Kerbel, N.A. Gentile, O. Jones, D. Munro, S. Pollaine, T.R. Dittrich and S. W. Haan, “Three-dimensional HYDRA 10. J.L. Linsky and B.E. Wood, “The alpha Centauri Line of Sight: D/H Simulations of National Ignition Facility Targets”, Phys. Plasmas, 8, pp. Ratio, Physical Properties of Local Interstellar Gas, and Measurement 2275-2280, 2001. of Heated Hydrogen (The 'Hydrogen Wall') Near the Heliopause”, Astrophys. J., 463, pp. 254-270, 1996. 20. H. F. Robey, et al., “Shock timing experiments on the National Ignition Facility: Initial results and comparison with simulation,” Phys. Plasmas, 11. B.E. Wood, J.L. Linsky, G. Hébrard, G.M. Williger, W.H. Moos and 19, 042706, 2012. W.P. Blair, “Two New Low Galactic D/H Measurements from the Far Ultraviolet Spectroscopic Explorer”, Astrophys. J., 609, pp. 838-853, 21. D. S. Clark, C. R. Weber, V. A. Smalyuk, H. F. Robey, A. L. Kritcher, J. 2004. L. Milovich, and J. D. Salmonson, “Mitigating the Impact of Hohlraum Asymmetries in National Ignition Facility Implosions Using Capsule 12. H. Kruger, P. Strub, E. Grun, and V.J. Sterken, “Sixteen Years of Shims”, Phys. Plasmas, 23, 072707, 2016. Ulysses Interstellar Dust Measurements in the Solar System: I. Mass

Received 9 May 2018 Approved 25 September 2018

JBIS Vol 71 No.4 April 2018 139 JBIS VOLUME 71 2018 PAGES 140–150

A SCIENCE-DRIVEN MISSION CONCEPT TO AN EXOPLANET

STACY WEINSTEIN-WEISS1, MARC RAYMAN1, SLAVA TURYSHEV1, ABHIJIT BISWASS1, INSOO JUN1, HOPPY PRICE1, ERIC MAMAJEK1, JOHN CALLAS1, TIM MCELRATH1, DAVE WOERNER1, JOHN BROPHY1, MIKE SHAO1, LEON ALKALAI1, NITIN ARORA1, LES JOHNSON2, MERAV OPHER3, SETH REDFIELD4, RALPH MCNUTT5, CAROL STOKER6, JENNIFER BLANK6, DOUGLAS CALDWELL7, LOUIS FRIEDMAN8, ROBERT FRISBEE8, GARY BENNETT8 1Jet Propulsion Laboratory, California Institute of Technology, 4800 Oak Grove Drive, Pasadena, CA 91109; 2NASA Marshall Space Flight Center, Redstone Arsenal Huntsville, AL 35812; 3Boston University, 1 Silber Way, Boston, MA 02215; 4Wesleyan University, 45 Wyllys Ave, Middletown, CT 0645; 5Applied Physics Laboratory, Johns Hopkins University, 11100 Johns Hopkins Rd, Laurel, MD 20723; 6NASA Ames Research Center, Moffett Blvd, Mountain View, CA 94035; 7SETI Institute, 189 N Bernardo Ave, Mountain View, CA 94043, USA; 8Consultant. email [email protected]

A concept for a science-driven robotic mission to an exoplanet was developed based on key mission and science requirements designed to address the question: “What makes a flight mission to an exoplanet compelling, in terms of science return, compared to what will be learned in the next few decades with large near-Earth telescopes or other remote sensing techniques such as a telescope at the Solar Gravity Lens Focus?” By thinking systematically through mission and science goals as well as objectives, key requirements were developed that would drive technology developments in all necessary aspects, not just on propulsion. One of the key mission science objectives was to confirm and characterize life. The team concluded that a direct confirmation of life would require in situ observations and measurements that cannot be performed on a fast (0.1c) flyby; thus, the mission would require a method to slow down, orbit or send a probe to the exoplanet’s surface. This capability drives a trade between interstellar travel velocity, trip duration and propulsion architecture as well as a high level of onboard autonomy, including adaptive science data collection, on-board data processing and analysis. This paper describes the mission concept, the key requirements and open trades.

Keywords: Interstellar, Exoplanet, Science, Mission, Spacecraft

1 INTRODUCTION ing a science-driven concept was important for determining what the mission needed to do upon exoplanet arrival. The Thirty years ago, the existence of planets orbiting other stars team debated whether a science-driven mission or a technolo- (exoplanets) was still unproven after centuries of specula- gy-driven mission would be the best mission concept to study, tion. Today, over 3,500 exoplanets have been discovered over since the answer has enormous ramifications for the science re- the past three decades, another thousand possible candidates turn and the mission technology requirements. In the end, the are awaiting confirmation, and the search techniques con- science-driven concept was chosen since it answers the ques- tinue to improve [1,2,3,4]. Recent identification of Habitable tion “what makes a mission to an exoplanet compelling?” and Zone [5] planets, such as Proxima Centauri b [6] and TRAP- it would best ensure the development of an extensible architec- PIST-1e [7], begs the question “When will a spacecraft be sent tural framework for the future. This choice should not exclude to investigate?” Thus, a study team was formed to develop a or diminish the value of precursor missions with other objec- science-driven mission concept to an exoplanet. The primary tives, such as to investigate the interstellar medium or validate mission objective was to confirm and characterize life at the required technologies – such precursor missions will likely be exoplanet, which is in-line with NASA’s strategic objectives required and their mission concept studies are encouraged. [8]. The team membership included experienced scientists and engineers from NASA centers, academia, independent institu- 2 SYSTEM DESIGN tions and consultants; members typically had participated in multiple space science missions. A key ground rule of the study The system design of an interstellar vehicle would be largely was to think “out of the box” and be creative, while simultane- driven by the propulsion system that is selected. More advanced ously prepared to defend innovative ideas with sound physics. systems, such as those using fusion or anti-matter, would be driven by the requirements for radiation shielding, magnetic The philosophy of developing a complete mission concept as plasma drive coils, and large radiators for dissipating waste heat opposed to focusing on one or two key technologies was used and heat from the shielding. A beamed energy sail would be to drive out mission-wide key requirements and trades. Choos- built around a thin film structure that might have completely different accommodation requirements for the science payload. A fission-pulse system would need to support a large and mas- This paper was presented at the Tennessee Valley Interstellar sive pusher plate that would drive the system design. Workshop 2017 Symposium, Huntsville, USA. Drawing from past work on interstellar mission studies

140 Vol 71 No.4 April 2018 JBIS A SCIENCE-DRIVEN MISSION CONCEPT TO AN EXOPLANET

• Increasing flight time • Descoping some of the payload • Switching to a closer target • Switching to fallback technologies

Starting up a serious effort for an interstellar program would require a systematic assessment of available technologies, possible but realistic near-term advancements, program- matic risk assessment, cost and schedule estimates, developing feasible system design concepts and identifying implementa- ble fallback options.

3 KEY MISSION REQUIREMENTS AND ASSUMPTIONS

A number of key mission requirements and assumptions were Fig.1 Nuclear Electric Propulsion Vehicle Concept. drafted early in the study and modified as the study unfolded. These notional requirements, assumptions and their rationale over more than half a century by many agencies, universities are listed below. and counries, a number of system options were considered in this study. Key system trades are shown in Table 1. Mission Duration: The threshold data shall come back within 70 years from launch. The most plausible vehicle approaches for a system that • Rationale: The threshold data must come back within the pro- could be built in the next 50 years would be nuclear electric fessional lifetime of someone born around launch; this person propulsion (NEP [7a, 7b, c]), and beamed sails; other systems can grow up learning about the mission and become inspired certainly could be built. An example of an NEP interstellar ve- by it, eventually joining the team and working to be ready to hicle design is depicted in Fig. 1. An example of a beamed en- interpret the data when it comes back. ergy sail vehicle design is shown in Fig. 2. If the spacecraft is travelling at a low fraction of the speed of light (0.1 – 0.3c), the exoplanet target must be no greater than The biggest system challenge for an interstellar vehicle is 15 LY of Earth (50 year travel time and ~20 years to collect and propulsion, and the degree of the challenge is a function of send the threshold data back to Earth). Note that this means flight time and ΔV. A mission that requires slowing down, stop- the first bit of data collected at the exoplanet will make it back ping, going into orbit, or landing on an extra-solar world more to Earth 15 years after exoplanet arrival. than doubles the propulsive requirements. The second biggest challenge is telecommunications. The optimum communica- Science Collection En Route: Science data shall be collected en tions system is probably one that minimizes the requirements route to the exoplanet. on the interstellar vehicle at the expense of a large Earth or • Rationale: There should be a mission conducted during the near-Earth based infrastructure to receive that data from the flight to the exoplanet to keep the science community engaged. spacecraft. For missions with flight times greater than 50 years, Meaningful science return at least every decade, with fields new technologies would need to be developed to ensure system and particles data being collected continuously and relayed reliability over the design life. to Earth on a regular basis, would significantly contribute to modeling and understanding the interstellar medium. An interstellar vehicle would be an unprecedentedly difficult undertaking and descope options would need to be defined in Launch Date: The launch date shall be no later than July 16, order to address and mitigate development risk. Some exam- 2069. ples of potential descopes would be: • Rationale: U. S. Congressional language introduced by Repre- sentative John Culberson [9]. This target allows five decades of exoplanet characterization to feed the target selection process, technology development and verification as well as maturation of scientific techniques and instrumentation, particularly in the area of life detection and characterization.

Confirm and Characterize Life:The mission shall seek to confirm and characterize life. • Rationale: Per NASA’s strategic objectives: “Discover how the universe works, explore how it began and evolved, and search for life on planets around other stars” [9]. The expenditure to develop infrastructure and technology to explore an exoplanet would be significant. Proposed near- Earth telescopes and/or a mission to the Solar Gravity Lens should be able to collect spectra that would achieve most of the exoplanet characterization objectives that a typical reconnaissance space science mission can achieve today [10,11,12]. In- formation from atmospheric composition (perhaps a next-generation LUVOIR or HabEx) to 1000 x 1000 pixel imaging of the Fig.2 Beamed Energy Sail Vehicle Concept. world (via a Solar Gravity Lens mission [13]) may be available

JBIS Vol 71 No.4 April 2018 141 STACY WEINSTEIN-WEISS et al

TABLE 1 Key system trades Flight Development Payload TRL Pros Cons Comments Time R isk Mass

Mission design

Fast flyby Possibly < 100 yrs 2 lower Minimum ΔV Encounter time is too short requirement

Braking at target > 100 yrs 1 high Adequate encounter Twice the ΔV of flyby time

Propulsion

JIMO-class, Isp NEP - JIMO -class ~12,000 yrs 2 lower small long flight times ~10,000 s

NEP - advanced ~1,000 yrs 2 lower large Might fit on a single Isp ~ 150,000 SLS

Beamed energy sail Possibly 50 yrs 2 high very small May require vast infrastructure Ref. Starshot

Fission pulse Possibly 200 yrs 2 high large Ref. Orion proj.

Beamed power EP > 500 yrs 1 high large Might fit on a single May require vast infrastructure SLS

Ref. BIS Daedelus, Fusion Micro-pulse Possibly 50 yrs 1 very high large inertial confinement fusion

Fission Fragment Possibly 50 yrs 1 very high large

Bussard ramjet Possibly 25 yrs 0 extreme large Minimal propellant No credible concepts required

No practical concepts for storing Antimatter rocket Possibly 25 yrs 1 extreme large antimatter or directing thrust. Anti- matter production and storage

Telecom

Optical com 4 lower

Large aperture μ-wave 3 moderate Might integrate with Difficult to maintain shape a sail

Power

Radioisotope 6 low

Fission 4 moderate

Beamed 1 high

Antimatter 1 extreme

(as an aside, space-based interferometers would have integra- 1000x1000 pixels or to 1 pixel with promising bio-signature tion times of thousands of years and thus are not practical). spectra. While the threshold for making a strong case for life is Thus, in order to have a compelling case to procure and spend expected to mature in the future with continued research, the the resources, a science objective is required that far exceeds suggested threshold is a start. the team’s anticipated near-Earth remote sensing capabilities over the next century. Moreover, a key lesson learned from the 4 SCIENCE OBJECTIVES AND REQUIREMENTS Viking lander experience [14,15,16] is to use multiple, unam- biguous investigations to confirm life; today, the only way to do Many science objectives can be envisioned for a mission to an that is via in-situ sampling. exoplanet, and eventually it will fall to the Decadal Survey process and NASA to decide upon them. For this study, five main Key Assumptions: categories of science objectives were considered: Heliosphere • The team cannot be constrained to today’s technology. An Boundaries, the Interstellar Medium (ISM) and other Science example is the use of 3-D printing technology in space En Route, Astrosphere of the Target Star, The Solar System of to build new parts for that spacecraft that have worn out the Target Exoplanet and the Target Exoplanet. during the long journey. • The exoplanet target has been previously observed such Of note to the first three categories, Voyager data appears that there is a strong case for life to show that ISM’s influence reaches further than previously The target exoplanet should have already been character- believed; the Voyagers are still being influenced by the Sun. ized via spectroscopy and/or imaging and either resolved to It is thought that a spacecraft probably needs to travel to 500

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AU from the Sun before escaping the Sun’s influence [17]. This has only had sufficient free oxygen (a potential biosignature) belief has implications for the study of the Sun’s heliosphere for less than a billion years. So for the Earth, advanced intelli- boundaries as well as the target star’s astrosphere boundaries. gent life has only been detectable for a world with a measurable Travelling at 0.1c offers almost 30 years of valuable scientific bio-signature for ~100 parts in 1 billion (~1x10-7). Thus, as a data collection within 500 AU. Realistically, key ISM and heli- proxy for other exo-worlds, there is a very small likelihood of ospheric objectives need to be achieved on precursor missions finding advanced life and there is a very small number of can- to an exoplanet mission; for spacecraft health and safety pur- didate exo-worlds available within 15 LY. poses, it is necessary to understand the environment that the spacecraft would encounter in order to fold that into the design Another example of a pitfall of relying solely on remote to achieve high confidence of success. However, it is still valu- spectra: the potentially habitable planet Proxima Centauri b able and it is a key baseline mission requirement, to conduct [6] is close to its star; the planet could have been (or still is) meaningful ISM investigations during a mission to an exoplan- subject to a massive solar wind and left with a thick O2 atmos- et, such as understanding the small scale structure of the local phere and yet no possibility of life [19]. In fact, there are at ISM and how that compares to the Sun/ Solar nebula, the ISM least three mechanisms for abiotic production of O2 on plan- composition, imaging rouge planets and measuring galactic ets with different geological histories [20,21]. Thus, spectral cosmic rays and short path-length emissions. In addition, oth- signatures alone are inadequate and the mission should orbit, er investigations could take place en route involving the extra- preferably landing instrumentation to sample exoplanet mate- galactic/ reionization background, extragalactic parallaxes at rial to confirm life. The team also discovered that a fast flyby nanoarcseconds and tests of general relativity; however, these mission of an exoplanet may not offer a compelling science re- can be performed on precursor missions as well. turn compared to near-Earth telescopes providing spectra and possibly imaging that could be achieved over the next decades. Science objectives involving the solar system of the target Flying by the target at 0.1c only gives ~100 hours in the tar- star are numerous and involve the typical basic reconnaissance/ get solar system, or ~1/2 of a planetary diameter per second. characterization objectives that missions in the solar system Even if quality data were to be captured on a fast flyby, sending have had – composition and mapping, atmospheres, moons, the data back when travelling at 17 AU/ day makes the prob- rings, dust, asteroids/ comets, refinements of size and mass, lem far more challenging. This a key finding of the study; to spin rates, etc. However, given that the mission’s main objective make a mission to an exoplanet scientifically compelling, the is study of one exoplanet in the star’s system, and there is a re- flyby must slow down – a fast flyby (0.1 c) is not scientifically quirement to return mission data within 70 years from launch, compelling. This finding had enormous ramifications on the it is not clear how many other exoplanets can be well-charac- propulsion design. terized given orbital mechanics and the mission duration. 5 TARGET CHARACTERIZATION AND SELECTION Science objectives involving the target exoplanet can include many of the basic categories; an orbiting mission can resolve Currently, 53 stars and nine brown dwarfs are known within rivers, forest, deserts, and oceans. A key mission requirement 15 LY of the Sun. It is remarkable that three brown dwarfs have is to confirm and characterize life. Returning to the question been discovered within 8 LY just within the past few years us- of what could be deduced via future near-Earth telescopes or ing the WISE infrared sky survey [22,23]. Of the stars iden- a mission to the Solar Gravity Lens versus a mission to an ex- tified within 15 LY, only nine are FGK-type stars that could oplanet: at a minimum, biosignatures of life must have been be roughly considered “Sun-like.” In the quest for optimizing detected at this exoplanet before choosing it as a target. Since observations for detecting biosignatures from potential life on this data will have been collected via remote sensing from long small rock-dominated exoplanets with surface liquid water, as- distances, these biosignatures would be spectral (disequilibri- tronomers have calculated theoretical limits of maximum and um components in the atmosphere such as O2, photosynthesis minimum stellar flux for stars of varying luminosity and/or ef- – red edge of vegetation), spatially-resolved images (structures, fective temperature appropriate for planets similar to Earth in cities, lights turning on and off, large-scale land modification), size and atmosphere (the so-called “habitable zone”). NASA is or electromagnetic (radio or optical signals). currently interested in studying the icy moons of Jupiter (Eu- ropa) and Saturn (Enceladus) due to their potential for harbor- Because these biosignatures are neither confirmation of life, ing life in liquid oceans under the ice. These types of worlds nor characterization of life, a mission to the exoplanet would be should not be excluded from exoplanet mission target consid- required to confirm and characterize its life. While a Solar Grav- eration in the future if the ability to put together a strong case ity Lens mission may be able to confirm existence of life (e.g. for life is developed. lights turning on and off, informational radio signals), it would not be able to characterize that life. Detection of atmospheric Today there is also a limitation on detecting and character- disequilibrium does not necessarily equate to a biomarker [18]. izing exoplanets that can be seen from Earth – either due to Atmospheric spectroscopy from afar can identify potential bi- the techniques currently in use, which can improve in the fu- osignatures better than imaging; however, a mission must go ture or due to the basic viewing geometry, period and radius of to the exoplanet to unambiguously confirm life with in-situ the planet, which cannot be improved. For example, the transit sampling, ideally with multiple, independent investigations. method, which has yielded thousands of planets on close-in or- A Solar Gravity Lens mission with 10 km imaging resolution bits around stars, is only sensitive to finding ~0.5% of planets could plausibly detect artificial illumination on the exoplanet, orbiting in the habitable zones (~1 au) of G stars due to the if present. However, the exoplanet may be a world where there random orientation of orbits among target stars; therefore, this is not yet advanced intelligent life to produce electric light. In- particular method is likely to yield few such planets among the telligent life capable of producing lights, radio signals, struc- nearest Sun-like stars. The study considered what future near- tures, etc. only recently appeared on the Earth and so there is Earth missions could do to find targets for a future mission to a low chance of finding life in that state. Those technologies go to another star. Precision radial velocity methods are cur- have only existed on Earth for about 100 years; the atmosphere rently challenged at the ~0.5-1 m/s level due to stellar noise

JBIS Vol 71 No.4 April 2018 143 STACY WEINSTEIN-WEISS et al sources and progress is being made to improve the method to astrosphere boundaries and the ISM would be very similar and the ~10s cm/s level needed for detecting Earth’s around Sun- likely include magnetometers, plasma detectors, cosmic ray like stars. Microarcsecond astrometry, with proposed large detectors, Lyman alpha detectors, radio detectors (for plasma space telescopes like LUVOIR or HabEx, may be able to survey density), dust detectors and ion/ electron direction and veloc- nearby Sun-like stars for ~Earth-mass planets on ~1 AU orbits. ity detectors. For short path length emissions, ultra-violet and The key technology elements for the LUVOIR High Definition X-ray spectrometers would be needed. An IR imager would imager to enable this level of astrometry is part of the baseline be useful for rogue planets. Finally, imaging the spacecraft LUVOIR instrument design [23]. Since G-like stars offer the shield would be useful to monitor damage over time from dust best prospects for finding life, given the current understanding and debris impacts. Simply monitoring the velocity over time of how to detect it, there are very few candidates within 10 LY through the ISM would yield data about the changes in proof Earth; thus, the mission requirements had to be pushed out gress due to the ISM. to 15 LY. If Habitable Zone planets around Alpha Centauri A or B are detected in the future and meet the selection criteria at Inside the target star’s solar system, a variety of cameras the time, they could make great candidate targets and are only ranging from narrow angle to wide angle with various resolu- ~4 LY from Earth; however, the team did not feel it prudent to tions would be desirable to look at distant planets and Moons; count on the existence of those planets today. the narrow angle camera could also be used to navigate closer in (see Section 9). Infrared and ultraviolet cameras and spec- In addition, the study considered what these future missions trometers with a variety of spectral ranges, in addition to visi- could reveal concerning apriori knowledge. LUVOIR [23a] ble and mass spectrometers, would be useful for thermal char- should be able to pin down the location of the exoplanet to with- acterization and compositional characterization. Also, at the in a few exoplanet radii, and the Solar Gravity Lens mission [23b] exoplanet itself, having one or more landers or probes with life should be able to pin it down to < 10 km. Either one of these detection experiments, a metrology station (wind, temperature, techniques should be good enough to plan a mission around. atmospheric density and composition) and a suite of cameras and spectrometers as described above, would be required. Target Selection Criteria The technology exists today, in limited scale, to use onboard It is understood that exoplanet characterization techniques and autonomy to allow instruments to detect targets of opportuni- the understanding of what constitutes a biomarker will expand ty (for example, dust devils on Mars, or a hurricane forming and improve in the future; therefore, the target selection crite- in the ocean on Earth). It is expected that future version of ria will adapt and change in the future. In fact, multiple papers this technology will be running onboard the spacecraft with a are in work/ recently published describing biosignatures [24]. priority scheme in order to increase the scientific return from However, based on limited knowledge today, the target selec- the mission. The spacecraft will also need this capability in tion criteria are: order to select a landing site and design a lander descent profile based on atmospheric and gravity data collected and assessed • Exoplanets that are in their star’s Habitable Zone autonomously.

• Exoplanets with masses > the mass of Mars (this would 7 COMMUNICATION indicate rocky planets with a decent chance for the existence of an atmosphere) All options considered (radio, optical) are constrained by the speed of light; the first bit of data returned from the exoplan- • Exoplanets that experience roughly the same solar radia- et will still take 15 years to travel back to Earth from 15 LY. tion as Earth Lasercomm was chosen for this study since the spacecraft will already have a camera on board for science. The OPALS la- • Detection of a biosignature from the exoplanet plus at sercomm system was recently tested aboard the International least 1 pixel image of the exoplanet (ideally 1000 x 1000 Space Station (ISS) and demonstrated significant increases in pixel image) transmit time and data volume [25]. However, a key mission consideration is the energy required onboard the spacecraft and • The current age and estimated lifetime of the star should it was shown that current technology - an OPALS-like system be such that life will have had a chance to form with 40 m light bucket receivers on Earth – would require over 100 kW of power to operate. A different approach was chosen – Th e current thinking is that the star should be at least assuming large aperture diameter transmitters and collectors to > 1 billion years old (and preferably older) increase net gain of the link.

• The exoplanet’s star should be close to a G2V Class (the For ground-based receiving of optical signal losses due to at- Sun) mospheric transmission of the Doppler shifted laser wavelength and irrecoverable atmospheric turbulence, induced aberration For instance, if Alpha Centauri A, which is estimated to be losses can be severe. On the other hand, space based receivers a six billion year old type G2V star, had a rocky planet in its will be free from atmospheric transmission and turbulence loss- habitable zone, it would be a great candidate to make it through es but they will need to be equipped with autonomy to search all the target selection criteria. for and acquire the laser link. In this study, the team aggressively chooses a 100-m diameter receiving aperture. 6 INSTRUMENTATION For any laser wavelength increasing transmitter diameter The instrumentation required for this mission would follow the results in higher far-field gain but also results in a narrower science and mission requirements set. In general, though, the laser angular beam-width needing tight pointing control. For instrumentation requirements for the heliosphere boundaries, this study, a telescope with Hubble Space Telescope (HST)

144 Vol 71 No.4 April 2018 JBIS A SCIENCE-DRIVEN MISSION CONCEPT TO AN EXOPLANET

TABLE 2 Notional Link Design for Interstellar Lasercomm

Average Laser 36.02 dB-W 4000 W 4 kilowatt average power 840 nm laser Power Transmitter Gain 138.17 dB 240 cm HST equivalent aperture diameter

Transmitter -6.21 dB Optical transmission and pointing losses Efficiency Space Loss -486.54 dB 15 ly Free-space loss from 15 lightyears

Receiver Gain 171.28 dB 100 m 100m collecting aperture in space Receiver Efficiency -10.56 dB Optical and implementation loss w/3-dB margin

Received Power -157.84 dB-W Photon Flux 28.42 dB-ph 695 ph/s Received photon flux in photons per second

like aperture and pointing control was adopted. The pointing protons and solar protons: surface charging from plasma elec- accuracy of the Hubble telescope is approximately 35 nano- trons and ions, internal charging from high-energy trapped radians (nrad) [26]. Assuming a laser beam-width of 10× the electrons, single event upsets from trapped protons, solar pro- pointing accuracy, as a rule of thumb, a 350 nrad beam can be tons, or galactic cosmic rays, structure damage and/or elec- transmitted. Given the HST-class aperture diameter of 2.4-m, trostatic discharge from dust or micrometeorites, and material a transmitted laser wavelength of 840 nm results in the beam- degradation from atomic oxygen and UV. Of these, the items width that can be accurately controlled. that are most concerning for a mission to an exoplanet are dust and galactic cosmic rays. Models by Weingartner and Drain Table 2 summarizes a notional link design for a 840 nm la- [27] and Hoang [28] show there could be up to 0.5 mm ero- ser with 4000 W average power transmitted from a 240-cm di- sion at velocities of 0.2c; this is an important consideration for ameter telescope. From interstellar ranges the sun is used as a design and especially for concepts involving ultra-thin light pointing reference to point the laser back at the Earth receiver. sails. A few mitigation approaches have been suggested, such The Earth receiver is a 100-m diameter space-based collecting as electric deflection or radiation pressure deflection [28]. In aperture with photon counting detectors. The laser is pulsed the ISM, interstellar galactic cosmic rays dominate the radi- with low duty cycle. ation concerns. However, a big unknown is the environment in the target star system, especially for high-energy radiation The laser communication signal will be discriminated from that is important for spacecraft design. X-ray observations of additive noise by relying on a combination of spectral and tem- the target host star could be used as a proxy for high-energy poral filtering. The spectral filters with a narrow linewidth will particle environment estimate although it will not be a direct be tunable to cover the Doppler uncertainty. Scanning the Dop- measurement of energy spectra [29]. pler uncertainty with dwell times sufficiently long to temporal- ly synchronize to the known pulse position modulation (PPM) 9 PROPULSION of the laser transmitter will be detect signal in the presence of noise. Active spectral and temporal tuning will lock the receiv- Since there is currently no existing propulsion technology that er to the signal and reject noise. The viability of this concept can achieve 0.1c, the team reviewed various propulsion op- as a means of communication in the presence of background tions in order to determine which candidates were the most and stray light noise needs further study. With stable pointing promising for technology readiness in the next five decades, of the 100 m aperture, 100’s of bits/second of data-rate can be assuming a robust and focused technology development. Op- received based on a 1-2 bits/photon link capacity. If the above tions considered for this study included matter/ antimatter an- link design were scaled to 4.37 Ly (Alpha Centauri), the pho- nihilation [30,31,32], beamed momentum [33,34], the Bussard ton flux would increase to ~ 8000 photons/sec supporting a few ramjet [35], Daedalus-style fusion [36], and fission fragment kilobits of data-rate. [37,38]. NEP was discarded since it did not meet the mission duration requirements established in Section 3. The require- While this indicates the viability of laser links from inter- ment for confirmation/ characterization of life resulting in the stellar distances, developing lasers that can survive the 50-70 need to orbit and land on the exoplanet is a huge challenge in year journey and operate reliably poses a formidable chal- the propulsion arena. lenge. 100-m apertures in space would also be non-trivial. The possibility of using gravitational lensing was explored and can From a propellant energy density standpoint, matter/ anti- offer huge gains except for the fact that the receiver would be matter is the top choice (Fig. 3) [39]. However, it is not a very nearly 550 AU from the sun and aligning the transmitter, grav- efficient system, decaying or radiating its energy before ther- itational lens and receiver poses a problem. malization [40,41,42] and thus it does not compare favorably with fusion or fission. In addition, only ~10 ng/ year of anti- 8 ENVIRONMENTS matter is produced versus the millions of metric tons required for a rendezvous mission that meets the stated requirements; it Understanding the space environment is key to spacecraft de- is unlikely that such large-scale production will be achievable sign. Concerns are radiation damage from trapped electrons, in the next five decades [43]. Special handling considerations,

JBIS Vol 71 No.4 April 2018 145 STACY WEINSTEIN-WEISS et al

CRUISE VELOCITY PROPULSION ALPHA CENTAURI 4.3 LY 40 LIGHT-YEARS CAPABILITY Matter-Antimatter Vision Mission Beamed Energy, Fusion Ramjet Fusion (Daedalus) INTERSTELLAR Fission Fragment

OORT CLOUD (6,000 AU-1 LY)

Advanced NEP

KUIPER BELT Saturn V (40-1,00O AU) Fission Thermal (NERVA) Electromagnetic HELIOPAUSE (100 AU)) Catapault GRAVITATIONAL LENS (550 AU))

TRIP TIME (YEARS)

Fig.3 Interstellar and Precursor Mission Cruise Velocity and Propulsion Requirements [44] such as magnetic levitation in ultra-high vacuum in a system missions such as achieving Pluto orbit in under four years, a that can never fail, are another challenge. mission to the Solar Gravity Lens in under 15 years, or sending 100 metric tons of cargo to Jupiter’s orbit in a year [48]. The Bussard fusion ramjet was considered and discarded because of its large number of technological issues. These in- 10 NAVIGATION clude the design of the electromagnet scoop, the need to collect interstellar deuterium for fusion (since pure hydrogen fusion Navigation for a mission to an exoplanet has to be autono- may never be achieved by humans), significant energy losses mous and on-board by nature in order to be useful since one- in the fusion reaction and the need for a “drag free” scoop, way light times are 4 – 15 years. Any ground-based navigation fusion reactor and electromagnetic nozzle [45]. Some of the would be purely forensic after about 500 AU. On-board, au- issues may have solutions; however, others may remain in- tonomous navigation was successfully used on Deep Impact tractable. For example, reactor energy losses; this issue could to hit the comet [49]. What has not yet been demonstrated is be overcome by laser power-beaming but then a laser sail may a fully autonomous on-board mission replanning system. To just as well be used without the mass overhead. elaborate, upon arrival at the exoplanet, the approach trajectory and target orbit parameters will need to be adjusted based on The three top choices for a potentially successful propul- new knowledge gained either about the exoplanet’s dynamics sion system development in the next five decades were a two- or the current performance capabilities of the spacecraft. In ad- stage laser sail [45a], fission fragment and fusion (Daedalus). dition, landing site selection would be based on data gathered All three of these proposed technologies have challenging and interpreted completely on-board the spacecraft rather than development problems associated with it; however, the team via teams of scientists on the ground. felt that the laser sail had the highest potential to be ready in five decades as well as the most potential payoff for the in- It was determined that today’s target knowledge is suffi- frastructure investment involved. The small (roughly 14-m cient to perform that mission. However, expected improve- square) Japanese Ikaros solar sail has flown in space [46,47] ments with LUVOIR and/or a mission to the Solar Gravity and the development challenges for this technology include Lens would certainly help. An extremely accurate time refer- deployment and control of ultra-large sails (> 100’s of meters ence will be required and thus the Deep Space Atomic Clock in diameter), developing ultra-lightweight sail material and the [50] appears sufficient for this type of mission. A consider- laser itself. The technology roadmap from today’s laser tech- ation is determining, from on-board optical navigation, how nology to what would be needed for a mission to an exoplanet close one is to the exoplanet; this can be achieved by observ- is daunting; yet, so are the development paths for the other two ing the planet’s motion over time via the onboard camera and potential propulsion candidates. In addition, there are political navigation system. ramifications concerning lasers of that power as well as space- based nuclear systems. It is expected that serious study be giv- 11 POWER en to all three options before committing to any single one. Advantages of a laser sail is that the propulsion source does Electrical power is arguably the most important subsystem on not have to be carried onboard, which allows the system to be any spacecraft because almost every other subsystem requires less massive and require less power to operate. Another nice electrical power. The basic elements of a spacecraft electric feature is that laser technology could continue to improve over power subsystem (EPS) are shown in Fig. 4. Essentially a pow- the course of the mission and those upgrades could be used by er source (e.g. nuclear or solar or power beaming) provides the the mission in flight. Like building the launch pads for early electrical power, which is conditioned through the power man- NASA rockets, a series of large (> terrawatts) lasers could also agement and distribution (PMAD) subsystem and other power be used in conjunction with laser-electric propulsion to enable processors to the “loads” (e.g. spacecraft instruments, comput-

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Power Power Management Load Source & Distribution Saturn

Jupiter

POWER SOURCES Energy • Power beaming Storage • Solar photovoltaic Mars • Solar dynamic • Radioisotope Power Sources (static or dynamic conversion) Earth • Nuclear reactors (static or dynamic conversion) Total area > 500 m2 ENERGY STORAGE • Batteries Two arrays • Fuel cells each 9m x 32m • Flywheels

Fig.4 The basic elements of a spacecraft electric power subsystem Fig.6 Relative solar array sizes as a function of distance from the (EPS.) Sun. ers, etc.). Energy storage is a way to save unused electrical able or having near-term availability (Fig. 4). Power beaming power for times when the spacecraft needs more power than offers the potential of having a low-mass electrical power sub- the power source can provide. For the proposed concept, the system because most of the mass from which the power would requirement is to provide 5 kilowatts (kW) for 70 years, which be beamed would be on Earth or in space. The two principal was based on 4 kW for the lasercomm system and 1 kW for all types of power beaming are lasers and microwaves. If the other spacecraft power needs. spacecraft were already using lasers for communication and navigation, power beaming would be a nice addition. Howev- There are five basic types of power sources currently avail- er, the spacecraft would need prime energy storage for times when the laser was down or the spacecraft fell off the beam.

Solar power has been used on most of the spacecraft that have been launched since the beginning of the space age. While the solar power has come from photovoltaic (i.e. “static”) conversion, solar power could also involve the use of “dynamic” power conversion (e.g. turbine-alternators or linear alternators), which offers the potential for higher thermal-to-electric conversion efficiencies. The 5 kW requirement is not an issue Solar radiation received at the for spacecraft operating in the inner Solar System; for exam- top of the Earth's atmosphere: ple, the ISS was designed for 110 kW of solar power. However, 1,368 watts per square meter. the major issue facing solar power subsystems is operating far from the Sun. With the solar flux falling off as the reciprocal of 1 Astronomical Unit (AU) the square of the distance from the Sun (Fig. 5) the solar array = 1.496 x 108 kilometers size would grow proportionately ( Fig. 6). (mean Sun-Earth distance) Currently there are two practical nuclear power sources: radioisotope and nuclear fission (at some future point, nuclear fusion may become an option for interstellar missions). A radioisotope power system (RPS) converts the heat from the natural decay of a radioisotope (e.g. plutonium-238 on U.S. space missions) to useful electrical power. To date, all of the RPS that have been flown by the United States have used ther- SOLAR ENERGY FLEX (EARTH = 1.0) moelectric elements for the thermal-to-electrical conversion, although research has been conducted on using other conversion systems (e.g. turbine-alternators and linear alternators). An RPS that uses thermoelectric conversion is called a radioisotope thermoelectric generator (RTG). MARS PLUTO EARTH The highest-power RTG flown by the U.S. is the General SATURN JUPITER URANUS NEPTUNE Purpose Heat Source-Radioisotope Thermoelectric Generator DISTANCE FROM SUN (ASTRONOMICAL UNITS (GPHS-RTG), which provided 300 kW at beginning of life (BOL) at about 6.8% conversion efficiency. Thus, almost 17 Fig.5 Relative solar energy flux as a distance from the Sun. GPHS-RTGs would be needed to provide 5 kW at BOL. How-

JBIS Vol 71 No.4 April 2018 147 STACY WEINSTEIN-WEISS et al ever, the natural decay of the plutonium-238 (87.7-year half- life) leads to a drop in power of at least 1.6% per year (plutonium-238 decay plus thermoelectric decay), which means there would be insufficient electrical power after 70 years. Note that the RTGs on the Voyager spacecraft have operated for over 40 years. If thermal control could be employed to keep the thermoelectric elements operating at the ideal temperature, it would theoretically be possible to have an RTG-powered mission if one started with about twice the number of BOL RTGs (about 34). A radioisotope decays away below detectable levels in about 10 to 15 half-lives; fortunately, a 70-year mission would be less than one half-life of Pu-238.

Luckily, there are higher-powered RPS concepts that could help reduce the mass needed to achieve 5 kW. The U.S. has sponsored work on kilowatt-class dynamic RPS using various power conversion technologies (Brayton, Rankine and Stir- ling), which offer the potential of achieving up to 30% conversion efficiencies. Thus, if one started with about a 10-kWe dynamic RPS, it would be possible to have about 5 kW available Fig.7 Artist’s Concept of a Nuclear Reactor Using Eight after 70 years (assuming no losses in the dynamic conversion 400-We Stirling Convertors. system). As an aside, it is worth noting that the Swan Falls dam in Idaho has turbines that have operated for 90 years. the beginning of the study. A prime example is the need to confirm and characterize life, which was not one of the original There are, of course, radioisotopes with longer half-lives objectives. Given anticipated advances in near-Earth telescopes than plutonium-238 [e.g. samarium-151 (96.6 years); nick- and/or a mission to the Solar Gravity Lens Focus in the next el-63 (100.1 years); silicon-32 (170 years); americium-241 50 years, it is expected that a considerable knowledge base on (432.2 years), etc.] that could be investigated to determine if the exoplanet would be amassed. The only compelling science they would alleviate the radioisotope fuel decay issue, recog- objective left that warrants the enormous investment in devel- nizing that the longer the half-life, the lower the specific ther- oping this mission would be to confirm and characterize life. mal power (watts thermal per gram of radioisotope fuel). Large space-based telescopes, some of which are currently A nuclear reactor can provide steady power through the ap- under study by NASA, will be required to characterize poten- propriate application of the control system (e.g. control rods or tial targets. Much work is still needed in the area of unambig- control drums). By removing the neutron-absorbing material uous life detection, particularly via remote sensing. Precursor in the control system, the power can be maintained at a steady ISM missions and/ or investigations are required to character- level as long as sufficient fuel (e.g. uranium-235) is available ize the environment the spacecraft would travel through and to fission. Significant tests would need to be run to ensure that test out required technologies, such as propulsion and long-du- a space-based reactor could operate for 70 years. The only re- ration on-board autonomous operations. actor flown by the U.S. to date (SNAP-10A launched in 1965) used uranium-235 as the fissile material (“fuel”); it functioned The single largest open trade is the propulsion technology for 43 days before an unrelated failure ended reactor opera- and associated sizing of that system. It is strongly recommend- tions. ed that a comprehensive trade study on propulsion technologies in the context of a mission concept be aggressively pursued, as Currently, NASA and DOE are investigating a reactor con- interstellar propulsion is bound to be a long-lead development cept that could provide 1 kW to 10 kW for missions up to 16+ effort. Just as the Solar Gravity Lens mission concept, based years (Fig. 7). In the past, the U.S. has studied a range of nucle- upon known physics applied in a novel way, has been recent- ar reactor concepts, from the kilowatt level to the multi-meg- ly proposed as a method to resolve an exoplanet, it is highly awatt level. desirable for a new, enabling propulsion technology to be proposed based on a novel application of known physical laws. With today's technology, the most reliable source of electri- While this mission was sized for targets within 15 LY from cal power for an interstellar spacecraft requiring 5 kW of con- Earth, the extensibility of the system concept for distances be- tinuous onboard power for 70 years would be an RTG. Howev- yond 15 LY was out of scope for this study. Since many more er, it is recognized that with current conversion efficiencies, this targets for extrasolar life exist beyond this distance, extending could require over 30 RTGs on the spacecraft. However, with the range to the target should be considered for future work. advances in space nuclear electric power technology, a nuclear reactor power system could become the lowest-mass nuclear Space power is an enabling technology and must be pursued power source with the fewest thermal and structural issues. aggressively. To power an interstellar spacecraft for 70 years at 5 kW, the following technologies should be pursued: 12 CONCLUSIONS • Improved conversion efficiencies (both for static and dy- A science-driven mission concept to an exoplanet has been namic conversion systems), including in-flight refreshing studied by a multi-center team of experienced engineers and scientists. Using a mission concept approach to the design, • Advanced radioisotope power systems (both RTGs and rather than focusing on just the key technology developments, dynamic RPS using Brayton or Rankine or Stirling con- resulted in findings and requirements that were not obvious at version cycles)

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• Long-lived, autonomous nuclear reactor power systems cepts should be examined before the one that is ultimately adopted for implementation is selected. Significant work re- It would be important to maintain the supply of enriched mains to be done on a science-driven mission concept to an uranium for future use in space reactors. exoplanet and the authors would like to encourage this work to continue. Significant work remains in on-board autonomous operations, including mission planning and system self-repair. Acknowledgements

The sizing of the lasercomm system is another area where The authors would like to thank Dave Gallagher and Gary open trades exist, including system sizing, laser frequency, la- Blackwood for supporting the formation of this study team. ser reliability over long durations and receiver locations. The research was carried out at the Jet Propulsion Laboratory, California Institute of Technology, under a contract with the Finally, it is acknowledged that many more mission con- National Aeronautics and Space Administration.

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35. R.W. Bussard, “Galactic Matter and Interstellar Flight”, Acta Inte4rstellar Sail: Starwisp Revisited”, AIAA, 2000. Astronautica, 6, pp.179–194, 1960. 35. R.W. Bussard, “Galactic Matter and Interstellar Flight”, Acta 36. A.R. Martin (ed.), “Project Daedalus—The Final Report of the BIS Astronautica, 6, pp.179–194, 1960. Starship Study”, JBIS, 1978. 36. A.R. Martin (ed.), “Project Daedalus—The Final Report of the BIS 37. B.G. Schnitzler, J.L. Jones and G.F. Chapline, “Fission Fragment Rocket Starship Study”, JBIS, 1978. Preliminary Feasibility Assessment”, Idaho National Engineering Lab., 37. B.G. Schnitzler, J.L. Jones and G.F. Chapline, “Fission Fragment Rocket Contract No. DEACO7-76IDO1570; also Lawrence Livermore National Preliminary Feasibility Assessment”, Idaho National Engineering Lab., Lab., Contract No. W-7405-ENG-88, 1989. Contract No. DEACO7-76IDO1570; also Lawrence Livermore National 38. R.L. Forward, “Radioisotope Sails for Deep Space Propulsion and Lab., Contract No. W-7405-ENG-88, 1989. Electrical Power”, AIAA, 1995. 38. R.L. Forward, “Radioisotope Sails for Deep Space Propulsion and 39. R.H. Frisbee, “Advanced Space Propulsion for the 21st Century”, J. Electrical Power”, AIAA, 1995. Propulsion and Power, 19, 2003. 39. R.H. Frisbee, “Advanced Space Propulsion for the 21st Century”, J. 40. J.L. Callas, "Technical Issues Associated with the Use of Antimatter for Propulsion and Power, 19, 2003. Spacecraft Propulsion", JPL, D-5877, 1988. 40. J.L. Callas, "Technical Issues Associated with the Use of Antimatter for 41. J.L. Callas, “The Application of Monte Carlo Modeling to Matter- Spacecraft Propulsion", JPL, D-5877, 1988. Antimatter Annihilation Propulsion Concepts,” JPL, D-6830 Rev. A, 41. J.L. Callas, “The Application of Monte Carlo Modeling to Matter- 2017, D-6830, 1989. Antimatter Annihilation Propulsion Concepts,” JPL, D-6830 Rev. A, 42. M.R. LaPointe, “Antiproton Powered Propulsion with Magnetically 2017, D-6830, 1989. Confined Plasma Engines”, J. Propulsion and Power, 7, pp. 749-759, 42. M.R. LaPointe, “Antiproton Powered Propulsion with Magnetically 1991. Confined Plasma Engines”, J. Propulsion and Power, 7, pp. 749-759, 43. R.H. Frisbee, “Optimization of Antimatter Rocket Performance”, AIAA, 1991. 2008. 43. R.H. Frisbee, “Optimization of Antimatter Rocket Performance”, AIAA, 44. R.H. Frisbee and S.D. Leifer, “Evaluation of Propulsion Options for 2008. Interstellar Missions”, 34th AIAA/ASME/SAE/ASEE Joint Propulsion 44. R.H. Frisbee and S.D. Leifer, “Evaluation of Propulsion Options for Conference & Exhibit, 1998. Interstellar Missions”, 34th AIAA/ASME/SAE/ASEE Joint Propulsion 45. B.N. Cassenti, “The Interstellar Ramjet”,AIAA , 2004. Conference & Exhibit, 1998. 45a. Frisbee, R. H., “Beamed-Momentum LightSails for Interstellar Missions: 45. B.N. Cassenti, “The Interstellar Ramjet”,AIAA , 2004. Mission Applications and Technology Requirements,” AIAA-2004-3567, 45a. Frisbee, R. H., “Beamed-Momentum LightSails for Interstellar Missions: 2004. Mission Applications and Technology Requirements,” AIAA-2004-3567, 46. E. Howell, “Ikaros: First Successful Solar Sail”, SPACE.COM, 2014, 2004. https://www.space.com/25800-ikaros-solar-sail.html [Last Accessed 8th 46. E. Howell, “Ikaros: First Successful Solar Sail”, SPACE.COM, 2014, September 2017] https://www.space.com/25800-ikaros-solar-sail.html [Last Accessed 8th 47. “Small Solar Power Sail Demonstrator "IKAROS" In Operation”, 2010 September 2017] http://global.jaxa.jp/projects/sat/ikaros/index.html [Last Accessed 8th 47. “Small Solar Power Sail Demonstrator "IKAROS" In Operation”, 2010 September 2017] http://global.jaxa.jp/projects/sat/ikaros/index.html [Last Accessed 8th 48. J. Brophy, “A Breakthrough Propulsion Architecture for Interstellar September 2017] Precursor Missions”, National Aeronautics and Space Administration, 48. J. Brophy, “A Breakthrough Propulsion Architecture for Interstellar 2017 https://www.nasa.gov/directorates/spacetech/niac/2017_Phase_I_ Precursor Missions”, National Aeronautics and Space Administration, Phase_II/Propulsion_Architecture_for_Interstellar_Precursor_Missions 2017 https://www.nasa.gov/directorates/spacetech/niac/2017_Phase_I_ [Last Accessed 8th September 2017] Phase_II/Propulsion_Architecture_for_Interstellar_Precursor_Missions 49. D. Kubitschek, et al., “The Challenges of Deep Impact Autonomous [Last Accessed 8th September 2017] Navigation”, J. Field Robotics, 24, pp.339–354, 2007. 49. D. Kubitschek, et al., “The Challenges of Deep Impact Autonomous 50. T. Ely, et al., “Expected Performance of the Deep Space Atomic Clock Navigation”, J. Field Robotics, 24, pp.339–354, 2007. Mission”, AAS/AIAA Space Flight Mechanics Meeting, Santa Fe, NM, 50. T. Ely, et al., “Expected Performance of the Deep Space Atomic Clock 2014. Mission”, AAS/AIAA Space Flight Mechanics Meeting, Santa Fe, NM, and Rockets, 22, pp. 345–350, 1985; also G. Landis, “Microwave Pushed 2014.

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150 Vol 71 No.4 April 2018 JBIS JBIS VOLUME 71 2018 PAGES 151–155

CONTACT WITH ALIEN BIOMES: Possible Biochemical Incompatibilities

KENNETH ROY and CATHERINE SMITH Tennessee Valley Interstellar Workshop, Oak Ridge, TN 37830, USA email [email protected]

Much effort has been expended on the search for habitable planets. One potential goal of this search is to find a second Earth, a planet that will eventually become another home for Humanity. While it is possible that lifeless, but habitable, planets exist, a more likely scenario is that a habitable planet has oxygen in its atmosphere because life of some sort is releasing it as a waste product. Earth life may not be compatible with alien life and attempts to colonize such worlds may not end well. Earth life is generally constructed of 21 specific amino acids. These amino acids are synthesized within the cells of living plants, animals, bacteria, archaea and fungi. Earth DNA systems are tailored to direct the construction of proteins from these amino acids. Yet, there are some 300 naturally occurring and an estimated 3000 plus possible amino acids that could exist. Independently evolved life on a distant planet will probably utilize some, but not all, of the 21 amino acids utilized by Earth life and will likely use some amino acids not used by Earth life. Ingesting (and perhaps even slight contact with) alien life could disrupt the function of Earth cells. Perhaps when humanity ventures to other solar systems, it will not be to colonize living alien worlds, but rather to find suitable sterile planets that can be terraformed into a second Earth inhabited solely by life transplanted from Earth.

Keywords: Alien life, Amino acids, Terraforming, Space Colonization

1 INTRODUCTION

It is currently believed that life first appeared on Earth over CARBOXYL four billion years ago [1]. There is some debate as to wheth- ACID er this life arose independently at that moment, evolving over GROUP time into what is seen today or if it arrived from some other source, such as Mars or further out. The latter theory is known as Panspermia [2]. If life arose on Earth independently, then it AMINO GROUP likely utilizes building blocks and molecules differently from alien life that arose on a similar world and evolved along parallel tracks. If the two forms of life interact, either through colo- Fig.1 A depiction of a typical amino acid. nization or exploration, then contact with the novel molecular building blocks and unique organic processes may result in tions delivered via messenger RNA (mRNA). perturbations in the delicate machinery within the organism and its cells. Alternatively, if life spread throughout the uni- These instructions originate in the nucleus of the cell and verse from a single source, as the Panspermia Theory postu- then travel into the cytoplasm, where the ribosomes are lo- lates, then this may not be a problem, as both may use similar cated. The ribosome then reads these instructions much like or identical, molecular building blocks. a punch tape. Based on instructions contained in the mRNA the ribosome then joins specific amino acids together in a se- 2 PROTEINS, MOLECULES THAT MAKE LIFE POSSIBLE quence. Amino acids have the ability to connect to each other with strong peptide bonds, thus forming one large molecule. A Proteins are complex molecules that perform functions neces- chain of amino acids linked together with these peptide bonds sary for life as we know it to exist. They can be structural ele- is called a polypeptide chain. If the chain consists of 50 or more ments within the cell or function like nano-machines or parts amino acids it is then generally referred to as a protein. of nano-machines. Proteins can also serve as messengers within the organism as in the case of hormones. All proteins are The mammalian ribosome is a complex nano-machine with strings of amino acids (Fig.1) that are assembled inside the cell a molecular weight of about 4.2 x 106 amu [3]. The ribosome by specialized molecular machines (ribosomes) using instruc- itself consists of approximately 80 polypeptides and proteins as well as 4 strands of ribosomal RNA (rRNA) that fit together in a very specific arrangement. The resulting ribosome is quite This paper was presented at the Tennessee Valley Interstellar stable and capable of many translation cycles. A typical mam- Workshop 2017 Symposium, Huntsville, USA. malian cell contains millions of these nano-machines. They as-

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TABLE 1 An assembly of amino acids. Image courtesy of Khan Academy: All Khan Academy content is available for free at www.khanacademy.org.

semble almost all of the billions of proteins that make up the no group of one to interact with the carboxyl acid group of an- typical Earth mammalian cell. other creating a peptide bond and releasing a water molecule. An example of a peptide bond formation is shown in Fig. 2. 3 AMINO ACIDS: THE BUILDING BLOCKS OF LIFE Interestingly enough, amino acids (except for glycine) can Each amino acid has an amino group, a carboxyl acid group be fabricated in the laboratory as one of two enantiomers, and a hydrogen atom connected to a central carbon atom. This called L- or D-amino acids, which are mirror images of each central carbon atom is also connected to an R complex, which other. However, all amino acids from Table 1 used for protein is different for each amino acid. Twenty of these amino acids synthesis by the ribosome (except for glycine), are L-amino ac- are shown in Table 1– the R complex is shaded. ids. It is noted that in the bacterial kingdom a few D-amino acids (methionine and leucine) have been found in proteins that The carboxyl acid group, under the correct pH conditions, are not synthesized by ribosomes [4]. The reason why Earth will tend to lose a proton. However, the amino group will ac- life chose L-amino acids over D-amino acids is unknown [5]. cept a proton, making the molecule overall neutral but highly polar and thus water soluble. The ribosome, following instruc- The amino acids used by living organisms on Earth are syn- tions contained in the mRNA, takes the appropriate amino acid thesized within the cells of living organisms through various and connects it to the previous amino acid by forcing the ami- pathways. A high percentage of amino acids making up any

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proper function of the cell machinery, the proper breakdown of nutrients to supply energy and production of amino acids. For example, two specialized enzymes are necessary to convert serine into selenocysteine [7].

The importance of enzymes to the proper function of the cell is illustrated by the development of Glyphosate. Glyphosate (N-(phosphonomethyl)glycine) is a broad-spectrum herbicide. It is an aminophosphonic analogue of the natural amino acid glycine and was first developed in 1950 by a Swiss chemist, Dr. Henri Martin. Glyphosate is water soluble and is absorbed by the plant mainly through foliage. It inhibits a specialized enzyme used in the synthesis of three aromatic amino acids – ty- rosine, tryptophan and phenylalanine – all fabricated through Fig.2 An illustration of a peptide bond formation. the Shikimate Pathway. When this enzyme is disabled the plant is unable to produce these amino acids. The plant then stops given Earth animal were originally synthesized in the cells of growing and soon dies [8]. Glyphosate, in theory, is harmless plants and other animals that were consumed as food. The to plants that do not use the Shikimate Pathway. digestive system’s mission is to process this food by breaking down food proteins into their constituent amino acids that are 5 ALIEN LIFE then released into the blood stream. The blood then carries these amino acids (along with oxygen, sugars and other nutri- “Two possibilities exist: Either we are alone in the ents) throughout the body, making them available to the cells. Universe or we are not. Amino acids are not stored in the body and are either used or Both are equally terrifying.” excreted if unnecessary. Arthur C. Clarke

Human cells can synthesize all but nine amino acids. These Astrobiology is the study of the origin and evolution of life nine are termed “essential” amino acids: histidine, isoleucine, beyond Earth as well as the effects of extraterrestrial environ- leucine, lysine, methionine, phenylalanine, threonine, tryp- ments on living things. Until alien life is discovered, it remains tophan and valine. If dietary intake is lacking in one of these a theoretical science. One reason to explore the solar system, essential amino acids, it is not possible for the ribosome to and hopefully find life, is to determine the validity of existing continue with polypeptide/protein production once the mRNA theories on life in the universe. Is it possible that alien life uses calls for the missing amino acid. This results in the cell actively silicon in place of carbon? Could it use carbon based chemical destroying other proteins in order to address the amino acid building blocks other than amino acids? Perhaps life is com- shortage. This usually results in more proteins being destroyed mon, arising wherever conditions are even remotely favorable. than produced. This process is sometimes referred to as mal- Then again, life could be rare, arising only when conditions are nutrition. perfect and then only occasionally. Currently, there is no way to answer any of these questions. Amino acids are used for more than just protein synthesis. They are used for synthesis of nitrogen compounds such as It is tempting, perhaps even reasonable, to assume that any creatine, purine, choline and pyrimidine, which can be broken planet similar to the primeval Earth in gravity, temperature down and used for energy. They are also necessary for the pro- and elementary composition will evolve a carbon-based life duction of a number of biologically active compounds. In short, system using amino acids as its basic building blocks. Life on these amino acids are vital for biological life as it is known [3]. Earth reveals that amino acids are useful as building blocks for assembling proteins and other biologically active molecules. Although there is some debate on the topic, there is another amino acid necessary for life. This 21st amino acid is seleno- As demonstrated by the Miller-Urey experiment, simple cysteine [6]; it has a structure similar to that of cysteine but amino acids result through natural process. This experiment with an atom of selenium taking the place of sulfur. Unlike the showed how simple chemicals, readily available on the pri- other amino acids, selenocysteine does not exist as a free-float- meval Earth, along with lightning, over time, could create nu- ing molecule waiting to be used. It is highly reactive and could merous organic chemicals, including some amino acids [9]. cause damage to the cell. Instead, selenium is stored in the Researchers like Christoph Adami, of the Keck Graduate In- cell as H2Se. When needed, the amino acid serine, which is a stitute of Applied Life Sciences in Claremont, California, and free-floating molecule inside the cell, is converted into seleno- his colleagues have looked at abiotic sources such as meteorites cysteine by removing the SH group and replacing it with a SeH and laboratory synthesis experiments and found that that sim- group. This newly created amino acid is then rapidly incorpo- ple amino acids like glycine and alanine are produced through rated into the polypeptide or protein chain by the ribosome. inorganic processes [10]. It appears that planets similar to the primeval Earth generate amino acids through inorganic pro- 4 ENZYMES AND DIGESTION cesses. At least, the simplest amino acids seem to be available in quantity, waiting for life to arise and make use of them. Enzymes are molecules that catalyze chemical reactions. Most enzymes are proteins and constructed in the cell like other However, that does not mean that evolution will automati- proteins. As catalysts, enzymes speed up chemical reactions cally result in life using the same set of 21 amino acids of which without being consumed. Some can help fold or even modify humans are comprised. There are over 300 naturally occurring polypeptides and proteins into the proper shape as they come amino acids to select from and many thousands of unnatural out of the ribosome. A sufficient set of enzymes is essential for amino acids that are chemically possible [3]. Researchers Davi-

JBIS Vol 71 No.4 April 2018 153 KENNETH ROY & CATHERINE SMITH la and McKay, in a paper published in Astrobiology, argue that quantities as a defense mechanism. Imagine if it were one of the chance plays a major role in the evolution of life and that life standard building blocks of an alien biosphere. Consumption that evolved on a planet identical to the primeval Earth would of such alien life forms could cause significant adverse effects. very likely use some amino acids found in Earth life but that it Even incidental contact might be problematic. is unlikely that they would use them all. It is also likely that they would use some amino acids not seen on Earth, except rarely All amino acids have “analogs’ in which one or two atoms and then usually in the laboratory [11]. have been switched out with a different element. Thialysine is an analog of lysine in which the second carbon of the R-group This scenario presents a problem. If humanity does reach has been replaced with a sulfur atom. Thialysine is not pro- the stars and finds a second Earth, identical in gravity and at- duced naturally but must be synthesized in the laboratory. It mospheric composition and even green with lush plant life un- is an amino acid with the same shape as lysine and has been der a G2 yellow dwarf star, then it is still unlikely that their ami- observed to replace lysine in protein synthesis. no acids would uniformly match Earth life amino acids. Weber and Miller concluded that if life were to arise independently on Thialysine and other amino acid analogs seem to be able to another planet, similar to primeval Earth, that about 75% of “fool” the ribosome into being used in protein formation re- the amino acids used would likely be the same found in Earth sulting in potentially defective proteins. In a recent study, it was life [12]. But, approximately 25% of amino acids used in alien demonstrated that rat cells exposed to analogs of arginine and life proteins would be foreign to Earth life. In addition, there is proline do incorporate them into proteins and that this results nothing special about the numbers 20 or 21. Perhaps on very in an observed loss of cell viability. The higher the concentra- challenging environments, larger numbers of amino acids are tion, the less viable the cells become eventually leading to cell necessary to afford life the flexibility it needs to survive. In death [16]. Such analogs to Earth amino acids, while rare on addition, the probability that the alien life set of amino acids Earth, could be common in alien life and could be one of their would include all of the essential amino acids needed by hu- standard amino acids. mans is small. Furthermore, potential problems are not limited to amino Additionally, there is perhaps a 50-50 chance that the ami- acids. Pheromones are chemicals secreted or excreted outside no acids present in the alien life would not be L-amino acids; the body to trigger a response in other members of the group. therefore, Earth life cells would be unable to utilize them prop- Most Earth organisms use this process to signal warnings, de- erly. Colonists could starve to death on a planet full of plants fensive behaviors or for reproductive purposes. Imagine if an and animals unless they brought their own food and crops or alien pheromone was capable of functioning as a psychoactive identified which nutritional supplements were required. Trying chemical or even a toxin to Earth life. A similar argument ap- to grow Earth crops on an alien world could also prove some- plies to plant pollen that can transport alien amino acids and what problematic. It might be possible, or it might be that the other chemicals through the air into the lungs of human visitors. soil biome would prove harmful, if not fatal, to Earth plants. Earth bacteria introduced into alien soil could find themselves 6 XENOBIOTICS defenseless against alien bacteria and might have trouble establishing nutrient cycles that they depend on. The opportunity Xenos means “stranger” in Greek. Xenobiotics are compounds for an ecological disaster for one, or both parties, cannot be that are strangers to Earth’s biosphere. This is typically used to dismissed. It is noted by the authors that this possibility might refer to the approximately 200,000 chemicals and substances help explain the Fermi Paradox. The malnutrition problem of that humanity has developed and now uses in large quantities consuming alien life forms has been noted in previous works within the last 100 years. Polychlorinated biphenyls (PCBs) are [13, 14]. one example of this. This chemical is very stable and has several useful properties. Unfortunately, when it enters the human But the problems of Earth life and alien life interacting go food chain, it and its decomposition products have proven to beyond nutrition and enter a potentially toxic situation. An ex- be very toxic to humans [17]. Other Xenobiotics include drugs, ample of this is the amino acid muscazone produced by the Eu- useful chemicals, insecticides, plastics, etc. Due to the hazard- ropean fly agaric mushroom (Amanita muscaria). Muscazone ous potential of new Xenobiotics, numerous agencies now re- is a psychoactive chemical compound that is toxic to animals view and approve them before they are introduced into Earth’s and people that ingest it. Ingestion causes visual damage, men- biosphere/environment. While far from perfect, these agencies tal confusion and memory loss [15]. The chemical makeup of are improving and may prevent future problems like the PCB muscazone is shown in Fig. 3. Mushrooms produce it in small debacle. However, these Xenobiotics are evaluated based on the potential harm to humans and Earth’s environment; alien biochemistries are not considered. The reverse is of course also true. If humans land on an alien world with an advanced civilization, they will find that it has its own large set of Xenobiotics that have not been evaluated against human biochemistry.

7 PANSPERMIA

The Panspermia Theory allows for the possibility that humans and aliens could interact with each other safely, at least on the cellular level. The Panspermia Theory proposes that life de- velops in one place and is then transported to other planets via bacteria riding on debris ejected from a planet during an impact event [2]. Ballistic, or interplanetary, panspermia pro- Fig.3 The chemical makeup of muscazone: a toxic amino acid. poses that life is distributed throughout a solar system by this

154 Vol 71 No.4 April 2018 JBIS CONTACT WITH ALIEN BIOMES: Possible Biochemical Incompatibilities mechanism [18]. This theory is supported by the discovery the alien amino acids could be toxic or even deadly if ingested. of meteorites found on Earth that originated from the planet Alien life would use many organic chemicals not seen by Earth Mars [19]. The Lithopanspermia hypothesis argues that life can life and some of those could be toxic or deadly. If humans en- spread beyond a solar system to nearby stars via the same, but counter an advanced technological civilization there is a ques- more energetic, impact events and that this is more likely dur- tion of xenobiotics in their environment, harmless to them but ing the early life of star when it is located in a star forming clus- perhaps not to humans. Even a long dead civilization might ter [20]. This theory allows for the possibility that life found on have left behind xenobiotics dangerous to humans. different planets within the same solar system or even planets around distant stars could be very similar in nature and could Ignoring the ethical question of colonizing a living alien perhaps be based on the same DNA information system and world, it would appear that establishing a colony on such a utilize identical sets of amino acids. This allows for the coloni- world is a potentially dangerous proposition. At the very least, zation of alien worlds by humans and Earth life without incom- there are unknown hazards present at the biochemical and cel- patibility issues -- or at least without significant compatibility lular level. Perhaps with advanced bio technology, the prop- issues as described in this paper. The Panspermia Theory could er nutritional supplements and/or drugs, it might be possible allow for taverns full of multiple alien species, often depicted in for humans to exist on an alien world; however, the long-term movies and science fiction novels. stability of such a venture is uncertain. Panspermia offers one possibility for humans being able to move into a living alien 8 CONCLUSION world with some degree of safety, at least on the cellular level.

With questions involving astrobiology, it must be understood Eventually, humans will venture among the stars to find a that there are many unknowns. If life evolves independently on new home for humanity. They may find that the best option a planet like Earth, it may use amino acids in the construction will be to select a suitable, barren, planet and terraform it with of proteins. However, life that evolves there may use D-ami- Earth life [21, 22, 23 and 24]. This would be a true second Earth no acids or may lack some of the essential amino acids needed for humanity. And perhaps as importantly, this would be a sec- by humans. This leads to the malnutrition problem. Humans ond home for the Earth life that has evolved with us and nur- could starve if they didn’t bring their own food or designer tured us throughout the millennia. Rather than seeking an ex- nutritional supplements. And it is not impossible that some of isting paradise, it may be that we should be prepared to build it.

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JBIS Vol 71 No.4 April 2018 155

DIARY FORTHCOMING LECTURES & MEETINGS OF THE BIS

APOLLO 8 – GETTING TO THE MOON BY DAVID BAKER 21 November 2018, 7.00pm VENUE: BIS, 27/29 South Lambeth Road, London SW8 1SZ SpaceFlight's editor relives his time working for the NASA Mission Planning & Analysis Division defining the mission's phasing and flight trajectories. He also casts an expert eye over the way mission planning evolved during the nine Moon-bound Apollo missions. THE INDIAN SPACE PROGRAMME 29 November 2018, 7.00pm VENUE: BIS, 27/29 South Lambeth Road, London SW8 1SZ BIS member Gurbir Singh talks about his 2017 book The Indian Space Programme – the story of India’s incredible journey from Third World country to First. CHRISTMAS GET-TOGETHER 2018 5 December 2018, 6.30pm VENUE: BIS, 27/29 South Lambeth Road, London, SW8 1SZ Join us for our usual relaxed evening of drinks and buffet food, along with our Christmas raffle. As is customary, this is a ticketed event to raise funds for the Society – and donations towards raffle prizes are always appreciated, too. Ticket price for members and their guests is £20. Guests are welcome, although we may have to limit the number of guest places if we sell too many tickets! APOLLO 8 – MEN TO THE MOON 18 December 2018, 7.00pm VENUE: BIS, 27/29 South Lambeth Road, London, SW8 1SZ Jerry Stone takes us on the next step in his series of 50th anniversary talks covering every Apollo mission up to and including Apollo 17 by looking back at Apollo 8's historic journey into lunar orbit – a triumphant end to an otherwise turbulent and tragic year. APOLLO MISSIONS: THE MECHANICS OF RENDEZVOUS & DOCKING BY DAVID BAKER 20 February 2018, 7.00pm VENUE: BIS, 27/29 South Lambeth Road, London, SW8 1SZ Starting with Apollo 9 launched on 3 March 1969, a key feature of the Apollo missions was the ability to rendezvous and dock in orbit – a capability that NASA had evolved over the preceding four years. SpaceFlight Editor David Baker describes the process in detail and casts an expert eye over the different options considered by mission planners in the run-up to the lunar landing missions. APOLLO 9 – RENDEZVOUS IN EARTH ORBIT 6 March 2018, 7.00pm VENUE: BIS, 27/29 South Lambeth Road, London, SW8 1SZ Jerry Stone continues his series of talks to celebrate the 50th anniversary of the Apollo missions with a uniquely personal take on the story of Apollo 9 – the first test of the full lunar landing package and only the second outing of the Lunar Module. Call for Papers RUSSIAN-SINO FORUM 1-2 June 2019, 9.30 am to 5pm (tbc) VENUE: BIS, 27/29 South Lambeth Road, London SW8 1SZ The BIS has now scheduled its 39th annual Russian-Sino Forum – one of the most popular and longest running events in the Society's history. Papers are invited. Watch this space for further details. Journal of the British Interplanetary Society

VOLUME 71 NO.4 APRIL 2018

PULSED MAGNETIC NOZZLE for Fusion Propulsion Jason Cassibry et al FISSION FRAGMENT ROCKET: Fuel Production and Structural Considerations Pauli Erik Laine FLYING ON A RAINBOW A Solar-Driven Diffractive Sailcraft Grover A. Swartzlander, Jr. EVALUATION OF THE HAZARD OF DUST IMPACTS on Interstellar Spacecraft Richard A. London and James T. Early A SCIENCE-DRIVEN MISSION CONCEPT to an Exoplanet Stacy Weinstein-Weiss et al CONTACT WITH ALIEN BIOMES: Possible Biochemical Incompatibilities Kenneth Roy and Catherine Smith

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ISSN 0007-084X PUBLICATION DATE: 23 NOVEMBER 2018