High Performance Solar Sails
and Related Reflecting Devices

High performance solar sails are light tension structures bearing space-manufactured, thin-film reflecting elements. They offer thrust-to-mass ratios 20 to 80 times those of proposed deployable sails. Development costs and risks appear modest. The low cost expected for sail production promises to make theme sails more cost-effective than solar electric propulsion for most missions of interest. Applications to near-Earth orbital transfers, deep space scientific missions (some unique), and non-terrestrial resource recovery are examined and found attractive. In the latter application, sails permit recovery of asteroidal resources with a very low initial investment. The promise of high performance, low cost, and great versatility recommend this system for further study.

Sporadic studies of solar sailing stretch back well over twenty years. For obvious reasons, nearly all serious studies have focused on launchable, deployable sails, made of necessity from comparatively rugged plastic film materials. The recent renewal of interest in solar sailing sprang from a JPL design study, which showed the feasibility of deployable sails of impressive performance.¹

Although these sails were attractive for a wide range of missions, a NASA review favored solar electric propulsion systems (SEPS) for further funding, owing to its better established technology base and lesser performance degradation in the outer solar system. Now, however, with the approach of the Shuttle era and plans for materials processing in space, it has become reasonable to consider the manufacture of sail materials in space. And this approach turns out to open a whole new game.

Space-manufactured thin-film materials promise sails with 20 to 80 times the performance of the best deployable sails; this would seem to justify a re-examination of solar sailing. To achieve this performance, methods must be devised to fabricate and handle delicate, unbacked, thin-film materials (the only other serious study of space manufacture of thin films that has come to the author's attention² produced foil about 500 times thicker than films of a thickness that the author has made, handled, and proposes making in space). In addition, the films must be incorporated into a structure that does not itself dominate the system’s mass.

These concerns are addressed in a systems design study³ that provides the foundation for this paper. This study also considers other issues related to the design, feasibility, and performance of the high performance solar sail (HPSS), including film materials, film properties, film element design and fabrication, sail assembly, forces on the sail, sail dynamics and attitude control, sail costs, and the lifetime of sails in the space environment.

High performance solar sails are truss structures built from tension members which support reflective panels assembled from vapor–deposited films 15 to 100 nanometers thick (see Fig. 1). Light pressure and payload inertial reaction provide axial tension, while slow rotation provides radial tension. HPSS's are expected to have high performance, low cost, and high reliability because of their low mass, ease of fabrication, and virtually passive mode of operation.

The core of the HPSS production system is a device which fabricates and mounts sheets of thin-film material. Film sheets are fabricated by depositing a sublimable material and then the reflective film onto a moving belt of metal foil, and then incising the film to separate it into sheets. The sheets are then mounted by gluing springy foil tabs to their corners, and freed by subliming away the intermediate layer through perforations in the foil or the film. This approach can apparently produce thin film sheets in a continuous stream while subjecting them to negligible loads. Two alternative approaches with different risks but similar results are proposed. Preliminary estimates suggest that a 3,000 kg device (inclusive of power supply) can produce some 3×10⁷ m² of film sheets per year. Another device assembles the sheets into triangular panels framed by tension members, which are accumulated for subsequent assembly to the sail structure. Together these devices make up a panel fabrication module, with a mass around 4,000 kg.

The sail's main tension structure is launched as a compact package, and then deployed from reels. Deployment takes place within the confines of a centrifugally-tensioned scaffolding structure incorporating six parallel beams. The main plane of the deployed sail structure is a hexagonal triangulated grid. A crane attached to the scaffolding conveys panels from the fabrication device to apertures in the grid, where they are hooked onto the structure. All fabrication and assembly operations may be accomplished without direct human intervention.

Air drag imposes an operational floor on solar sails between 700 to 900 km altitude. If sails are made below this altitude, they must be kept in orbit during manufacture by constant thrusting. Thrust requirements have not been estimated, but may be minimized by proper choice of sail attitude.

If sails are made above this altitude (prior to the establishment of a high-orbit station), all equipment must be designed for maintenance or replacement by teleoperator, owing to the radiation environment of intermediate altitude orbits.

After panel installation and release from the scaffolding, the sail becomes operational. The sail’s attitude to the sun then determines the direction and magnitude of its thrust. Since the sail spins, changes of attitude require a precessing torque. Tilting some of the panels produces torque for spin-rate control or slow precession; shifting the payload to an off-axis position in inertial space will produce torque for faster precession. in near-Earth maneuvers, the latter method would be used, permitting the sail to precess at rates around five radians per hour. In interplanetary flight, the first method can be used, and the sails may be left passive for weeks at a time.

Similar film fabrication devices can produce reflecting materials for other uses in space. Although the structure and film elements designed for sails would be poor for concentrating or directing sunlight (since their flatness has been sacrificed in various ways to save mass), thicker, flatter fins made by similar techniques and mounted under greater tension might serve. On a solar power satellite (SPS) incorporating concentrators, their use could save some hundreds of tons of plastic film mass, and night save substantially more mass in the structure if they required low mounting tensions commensurate with their small thicknesses. The cost per unit area would be somewhat greater than that for the sail, owing to the greater thickness required (100 to 500 nm?) and consequent higher cost of the materials and their evaporation.

The only system elements embodying substantially new technology are the film fabrication device and the film sheets themselves. The fabrication device has moving parts comparable in complexity to those of the beam-builders already under development for space use. Adequate control of the various substances handled will require careful design (baffles to confine vapors, bearings sealed against loose flakes of vapor-deposited materials, etc.). The device is designed to deposit and free the film sheets under conditions better than those commonly employed in the laboratory for producing unbacked films even thinner than those proposed for use in the HPSS. Several candidate film materials are fairly well characterized, and have strengths over 1,000 times what the present application demands.

The greatest uncertainties in the system – the minimum practical film thickness and the optimal film composition – have little effect on the system design. They affect the sail’s acceleration and maneuverability, but the proper choice of sail rotation rate negates effects on the sail’s structural design or major internal modes of vibration. They affect the design of the evaporators in the film fabrication device (and the sizing of their power supply and cooling system), but the design of the rest of the sail manufacturing system remains unaltered. Figure 2 indicates the range of uncertainty in the sail’s final mass and performance. Since the worst likely case yields excellent performance, and since these uncertainties are effectively decoupled from program risk, they may be greatly reduced by a very modest experimental program of film deposition and testing.

The balance of the system presents the typical problems of the design of a reliable, maintainable mechanical system. A preliminary estimate of the masses of the system elements to be developed is five tons. Applying a typical aerospace development cost of $20,000/kg to this mass yields a crude development cost estimate of $100 million. A substantially higher development cost could clearly be tolerated, if the system proves attractive enough.

The total mass of a system able to produce sails of 1 to 100 Newtons thrust (0.5 to 5 km diameter) is estimated to be around 13 tons (a smaller scaffolding with lower mass is apt to be built during development). Such a system could produce sails with thrusts totaling 200 N in a year’s time, and this capacity could be expanded in 200 N/yr increments by the addition of four– ton panel fabrication modules. Sails of greater thrust could be made by clustering smaller sails or by means of the construction of a larger scaffolding. The basic system, together with raw materials for sails totaling several hundred Newtons thrust can apparently fit within the mass and volume constraints of a single Shuttle launch.

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To be of practical interest, the HPSS must outperform its competition in a sufficient range of applications. At present, the competition is SEPS, using either ion engines or possibly mass drivers for conversion of electric to kinetic energy. The latter system is an efficient, omnivorous mass accelerator proposed for industrial-scale space propulsion.

Two performance parameters commonly compared among thrusters are specific impulse and thrust to mass ratio. Strictly speaking, the specific impulse of any sail is infinite, as it expends no mass. Still, sheet metal makes poor sails. In the spirit of amortization applied to durable goods, a performance measure related to specific impulse may be calculated by introducing a fictitious expenditure of 10% of the vehicle’s mass per year, and adding this to any mass actually expended. Sail thrust characteristics are also unusual as a sail varies its angle to the sun, its thrust varies in both magnitude and direction. For heliocentric trajectories, the useful component of thrust is frequently that perpendicular to the radius vector from the sun. On this basis, the useful thrust of a sail is about 38% of its maximum thrust.

Figure 3 plots the measure of "specific impulse" described above as a function of the (useful) thrust–to–mass ratio for sails of varying mass per unit area, at Earth’s distance from the sun, and SEPS with varying specific powers and exhaust velocities. As may be seen, the HPSS greatly outperforms state-of-the-art, multi-mission ion-engine SEPS, in both dimensions. An idealized (or mass driver) SEPS can exceed the sail’s thrust-to-mass ratio, but only at a low specific impulse, or with a very high specific power.

For industrial applications, the unit cost of transportation is of central concern. SEPS costs will depend on the delta–V of the mission, and on the costs of the reaction mass and the kinetic energy in the exhaust (the latter amounts to the cost of vehicle amortization). HPSS costs will depend on the cost of the sail, its distance from the sun, and the efficiency with which its thrust may be used. The following comparison will neglect other costs for both systems, and hence cannot be considered to represent total costs.

Figure 4 presents such a cost comparison. Table 1 summarizes the assumptions used, intended to represent an early era of space development. As may be seen, sail costs are lower than state-of-the-art SEPS costs across the board. The advanced, idealized, mission-optimized SEPS beats the more expensive sails at low enough delta-V, if a cheap enough source of reaction mass can be found. However, even the cost of mass–driver derived lunar materials has been estimated at over $1/kg.⁴

Figure 5 graphs comparative costs in the context of an SPS program. Table 2 lists the assumptions used. Once again, SEPS appears competitive only with low reaction mass costs, and at low delta–V’s.
 
 

Figure 4. Comparison of HPSS and SEPS costs.	Figure 5. Comparison of HPSS and SEPS costs in the context of an SPS program.
Table 1. Cost assumptions for Figure 4. General Assumptions: Discount rate, 10%. Cost of equipment purchase, $600/kg. Cost of transport to orbit, $700/kg. Only the cost of equipment, materials, and their transport to orbit is considered. HPSS: Cost of structure purchase, $600/kg. Cost of evaporator feedstock purchase neglected. One panel fabrication module used at capacity in a large scaffolding. ---The resulting cost is 12 to 26 cents/m2, depending on film thickness. SEPS: "State-of-the-art SEPS" curve represents a JPL multi-mission design6, costed as above. Other SEPS are assumed to have mission-optimized exhaust velocities, and to be able to supply kinetic energy to the exhaust for 25cents/kW-hr. This corresponds to a 70% efficient system with a mass of about 12 kg/kW (electric), costed as above. Curves are plotted for various assumed reaction mass costs.	Table 2. Cost assumptions for Figure 5. General Assumptions: Cost of transport to orbit, $100/kg. Other assumptions as in Table 1. HPSS: Cost of structure purchase, $200/kg. Cost of evaporator feedstock neglected. Many panel fabrication modules per scaffolding. $600/kg for initial module purchase, 80% learning curve applied to the cost of enough modules to provide 100,000 tons/yr transport capacity to geosynchronous orbit after 5 years production. ---The resulting cost is 1.5 to 3.5 cents/m2, depending on film thickness. SEPS: Same assumptions as in Table 1, but 2.5 cents/kW-hr (a commonly quoted target cost for SPS electrical energy delivered to the ground) is assumed as the cost of kinetic energy in the exhaust.

The HPSS can transfer payloads from low Earth orbits to geosynchronous orbit, from low inclination orbits to high, and so forth, with no expenditure of mass. Preliminary estimates suggest that a 10 N sail can transfer a 3 ton payload to geosynchronous orbit in less than 50 days. A 100 N sail could do the same with an entire shuttle payload. In the near term, air drag will keep sails from picking up payloads below an altitude in the range of 700 to 900 km. In the longer term, tethered satellite technology should permit sails to serve lower altitudes without dipping too low themselves.

Most of the light reflected from an HPSS spreads in a cone some tens of degrees across, and need never be directed at Earth’s night side. Still, some light will be scattered in all directions, and sails maneuvering at a few Earth radii will be visible over much of Earth’s night side for part of every orbit. Scattered light might become bright enough to affect astronomical programs involving faint objects. If so, the problem may be alleviated, even for heavy sail traffic, by arranging sail schedules so as to leave the skies completely free of sails most of the time (convoying). Eventually, instruments in space will more than make up for lost observing time on the ground.

At some altitudes, orbital debris would pose a substantial hazard for sails, if they were left uncontrolled. All significant orbiting objects are tracked, however, and hence may be avoided. Further, HPSS technology can be scaled down to produce small, cheap vehicles with unlimited delta–V capability. Such vehicles, with suitable payloads, would make fine orbital garbage trucks, allowing debris to be collected economically.

The HPSS can accomplish all of the great variety of missions proposed for deployable solar sails,’ and more. In general, its use will yield increased payload, reduced cost, reduced mission time, or all three.

• Such sails can serve as reusable interplanetary shuttles for delivery of orbiters, landers, penetration probes, etc., and can return samples from low planetary orbits to low Earth orbits.
• Their unlimited delta–V capability allows asteroid survey missions of indefinite length.
• Their high performance permits a rendezvous mission to Halley’s comet with a flight time well under a year, and permits pre–perhelion rendezvous missions, with subsequent sample return, to objects on parabolic trajectories.
• Their high performance permits not only fast out– of–the–ecliptic missions, but establishment of permanent solar polar observatories (see Figs. 6 (a) and (b)).
• With the aid of rocket stages, they can deliver large payloads into orbits around the outer planets with short mission times.
• They can perform 1.5 year flyby missions to Neptune and Pluto.

In lunar resource recovery scenarios, the HPSS can serve a variety of roles. It can transport equipment and supplies cheaply from low Earth orbits to low lunar orbit. In systems using mass drivers, it can help stationkeep mass catchers (perhaps increasing the range of feasible launch sites for achromatic trajectories), and can transport material from the collection point to a high Earth orbit for processing. In systems using rockets for lunar launch, it can transport materials from low lunar orbit to a high Earth orbit.

It seems more natural, however, to apply the low cost transportation capabilities of the HPSS where they can be used best: for asteroidal resource recovery. Certain Apollo objects should yield sail delta–V’s to geosynchronous orbit in the 5 km/second range, suggesting costs of $1-3/kg for low rates of mass recovery. These costs are comparable to those of lunar mass driver systems of much higher throughput,⁴ and would fall with increasing capacity. Since the resources returned are of comparable quality to those of the moon, these considerations alone make the HPSS/asteroid scenario attractive.

Historically, a great barrier to the use of non-terrestrial resources has appeared to be the high initial cost of the recovery systems. Nothing short of an SPS program has appeared to justify the expense of the development, construction, and emplacement of a lunar base, mass driver, mass catcher, and other system elements necessary for the lunar scenario. The HPSS, on the other hand, provides a reliable, low-cost, deep-space transportation capability well suited to operation without crew maintenance. With it available, the threshold to non-terrestrial resource recovery may apparently be crossed with a single Shuttle payload.

Figure 7 illustrates one approach to the surface mining of a small asteroid, based on a device which sweeps up loose surface matter and places it in a bag. Such a device may have many redundant sweeping heads, and seems unlikely to require human attention. A 200 ton sail-load, appropriate to a 100 N sail, may be swept up in under a month at a rate of one tenth kilogram per second. A few I millimeters thickness of loose surface material would suffice for many loads of this size, which may be returned with trip times on the order of a year. Two accessible asteroidal bodies with much loose material are already known: the moons of Mars. It would be ironic if they proved more attractive than our own.

These characteristics of HPSS/asteroid resource recovery systems open a range of non-SPS scenarios for space development. Demand for a few hundred tons of asteroidal material for radiation shielding could justify mining operations. Military demand for asteroidal steel to harden orbital installations could easily exceed 10,000 tons (or 100,000 tons, for that matter). Mass transport rates of this order of magnitude would drop the total amortized program cost per kilogram returned into the $2-20 range. Incremental costs for sail production would be low, and incremental costs for the use of existing sails would be almost negligible.

Many asteroids apparently contain a good grade of steel, with a typical cobalt content around 1% and a nickel content around 10%. Sail transportation costs from Apollo objects, at substantial traffic levels, should fall below 50 cents/kilogram. This may be compared with the market prices and world demands for cobalt (about $10/kg and 20,000 tons/year) and nickel (about $5/kg and 700,000 tons/year). If a suitably low-cost concentration or purification process can be found, and if return of materials through the atmosphere proves as inexpensive as expected,⁵ these metals might be sold on Earth. World markets are several billion dollars.

A large market might also be developed for foam steel. It should be easy to produce in space, and would require little refining of the raw material. Its unique properties might bring a price of several dollars per kilogram (roughly comparable to that of some wood products).

Preliminary estimates of the cost of mass– produced solar sails, using non-terrestrial feedstocks and a Shuttle–derived heavy lift launch vehicle for equipment transport, fall around 0.75 cents/m², yielding transportation costs around 6 cents/kg from a suitable asteroid. Since pig iron sells for over 20 cents/kg, and steel bars, plates, etc. for over 40 cents/kg, a non-terrestrial steel production industry is not out of the question. U.S. demand for pig iron is presently over 70 million tons per year; worldwide demand is almost 500 million tons per year. Cheap process heat, zero gravity, and accessible vacuum all make space attractive for steel processing. Steel and cheap energy as a basis for industrial development is an old story.

The HPSS is a high–performance, low-cost transportation system based on space–manufactured thin films. It is produced by a panel fabrication device comparable in complexity to a beam-builder, together with a low-mass scaffolding structure. Crude estimates suggest a very low development cost, hence actual costs are apt to prove modest.

Once developed, the HPSS can help probe the sun, near interstellar space, and points between. Its capabilities suggest use for transfer of payloads from low to high Earth orbits, and between orbital planes. Costs are low, as the sails themselves are inexpensive and expend no mass.

As a bonus, the HPSS can return asteroidal materials for use in near-Earth space with a small initial investment, and can return them in quantity for a very low cost. Applications to SPS programs are obvious, but these capabilities also open the door to an evolutionary expansion of non-terrestrial materials use in no way dependent on the SPS or on the near–term production of low-cost power systems in space.

In light of the unique capabilities, extreme versatility, low transportation costs, and modest development cost promised by the HPSS, this concept seems worthy of review by aerospace groups. If this promise holds up under closer scrutiny, work should begin to improve sail and sail fabrication facility design, to improve understanding of the proper role of the HPSS in space activities, and to initiate development of the sail fabrication technology itself.

The author would like to thank the National Science Foundation for its support during the greater part of this work, and the members of the JPL solar sail design teams: without them, this work would never have been done.

1. Friedman, L., et. al., "Solar Sailing-The Concept Made Realistic," 16th Aerospace Sciences Meeting, January, 1978, AIAA Paper No, 78-82.
2. Lippman, N. E., In-Space Fabrication of Thin-Film Structures, NASA CR-1969, Astro Research Corporation ARC-R-410, 1972.
3. Drexler, K. E., Design of a High-Performance Solar Sail System, M.I.T. Space Systems Laboratory Report 5-79, 1979.
4. Chilton, F., B. Hibbs, H. Kolm, G. K. O’Neill, and J. Phillips, "Mass Driver Applications," Space-Based Manufacturing from Nonterrestrial Materials, Vol. 57, Progress in Astronautics and Aeronautics, G. K. O’Neill, ed., 1977.
5. Gaffey, M. J., and T. B. McCord, "Mining Outer Space," Technology Review79(7):50, June, 1977.
6. Austin, R. E., R. E. Dod, and D. Grim, "Solar Electric Propulsion for the Halley’s Comet Rendezvous Mission: Foundation for Future Missions," 16^th Aerospace Sciences Meeting, January, 1978, AIAA Paper No. 78-83.

Q. Could you create a big enough solar sail to bring back an entire asteroid rather than small chunks of it?
A. It would have to be a very small asteroid to be practical--smaller than the ones we’re finding these days.

Q. What is the highest solar–system escape velocity a solar sail could achieve?
A. Maximum escape velocity is best attained by driving as close as possible to the Sun. The limit then depends on what temperature the film can stand. Using what I believe to be conservative laboratory data on unbacked thin film, it might be possible to achieve exit velocities from the solar system in excess of 1500 km/sec.

Q. Suppose you have a sail which is carrying no payload. How many solar g's acceleration would you get from the light pressure?
A. In the range of five to twenty.

Q. So the solar sail can carry many times the sail mass?
A. Yes. The small 10-newton-thrust sails I mentioned can transport about 40 tons/yr across a 5 km/sec delta-V. At about 12 tons/km² a 10–newton sail would mass about 15 tons.

Q. Have you considered the effect of the solar wind on the thin films? Doesn’t it cause considerable degradation?
A. I have considered that effect. The overall issue is what happens to the films on long-term exposure to the space environment. Micrometeoroids, for example, erode the area only very slowly, there’s no problem there. The solar wind particles cause sputtering; that is, the solar-wind atoms impact the sail at about 500 km/sec and knock atoms loose from the film. However, I have a good deal of data on sputtering both from estimates on transmission sputtering with thin films that were done for fusion work, and from measurements of erosion of the lunar surface by the solar wind. Based on these data, sputtering does not appear to be serious over time scales of a century or so.