DIRD Laser Lightcraft Nanosatellites

Laser Lightcraft Nanosatellites

Defense Intelligence Reference Document, Defense Futures. DIA-08-1011-001, 01 November 2010 (IcOD: 30 August 2010).

This product is one in a series of advanced technology reports produced in FY 2010 under the Defense Intelligence Agency Advanced Aerospace Weapon System Applications (AAWSA) Program.

Summary

Miniaturized satellites are spacecraft of unusually low mass and small size, usually under 500 kg in total mass. The term "minisatellite" refers to a spacecraft with a wet mass (including onboard propellant) of 100 kg to 500 kg. Microsatellite or "microsat" is a spacecraft with a wet mass of 10 kg to 100 kg. Nanosatellite or "nanosat" is a spacecraft with a wet mass below 10 kg. Picosatellite or "picosat" is a spacecraft with a wet mass of 0.1 kg to 1.0 kg. Picosats are also called sub-nanosats.

The primary reason for miniaturizing satellites is to reduce cost. Heavier satellites require larger launch vehicles of greater cost while smaller, lighter satellites require smaller and cheaper launch vehicles and can sometimes be launched in multiples or "piggyback", using excess capacity on larger launch vehicles. Miniaturized satellites allow for cheaper designs as well as ease of mass production. However, few satellites of any size other than communications constellations, where dozens of satellites are used to cover the globe, have been mass produced in practice.

Besides the cost issue, the main rationale for the use of miniaturized satellites is the opportunity to enable missions that a larger satellite cannot accomplish, such as:

• Constellations for low data rate communications.
• Using formations to gather data from multiple points.
• In-orbit inspection of larger satellites.

Many of these missions require numerous small spacecraft in a constellation or "swarm". These include orbital communications networks and swarms of small satellites to conduct remote sensing, and to provide unique perspectives on astronomical bodies of interest. For instance, 100 or more nanosats could be deployed from a mother ship to their final destination in space for deployment.

Provisions for orbital maneuvers as well as attitude control, multiple sensors, and instruments, and full autonomy will yield a highly capable miniaturized satellite. All onboard electronics will survive a total radiation dose rate of several hundred kilorads over a several year mission lifetime (at least 100 kilorads over two years). Nanosats developed for in-situ measurements will be spin-stabilized, and carry a complement of particles and fields instruments. Nanosats developed for remote sensing measurements or for surveillance and eavesdropping will be three-axis stabilized, and carry a complement of imaging and radio wave instruments. Autonomy both onboard the nanosats and at the ground stations will minimize the mission operational costs for tracking and managing a constellation.

To reduce overall mission cost, advanced technology components and a novel laser propulsion system will be used to make nanosats and their onboard instruments compact, lightweight, low power, low cost, and able to survive their radiation environment over a several year lifetime. Each nanosat will be manufactured and tested for a recurring cost not to exceed 500 thousand dollars. By producing a large quantity of nanosats for a given mission, the per-unit cost will be reduced to a small fraction of satellite procurements for traditional missions. Mission operation costs will be minimized by the incorporation of both onboard and ground autonomy and use of heuristic systems.

Chapter 1: Nanosatellite Technologies

Overview

Nanosats require technologies that radically reduce the mass and power of components without compromising performance. In addition to miniaturizing components, methods to integrate similar functions across subsystems are being evaluated. For example, all subsystem electronics, including instruments, could be integrated within the Command and Data Handling subsystem. Multifunctional solutions also offer significant savings over traditional approaches. Technology investments are required to develop or adapt components to accommodate the expected radiation environment. Simple, effective methods of thermal control are essential to keep the nanosat operational during extreme temperature variations. Autonomy is a critical technology that impacts every subsystem. Constellations with tens to thousands of nanosats must be highly autonomous to be practical. The nanosat ground system must be kept inexpensive, simple, and made inter-operable with other missions.

Propulsion

In the baseline mission, nanosat propulsion is needed for two distinct functions: first, each nanosat must raise its orbit apogee to the appropriate radius; second, it must reorient the axis of the spinning nanosat from the velocity direction, within the orbit plane, to its science mission attitude, perpendicular to the ecliptic plane. These maneuvers present challenging velocity change and attitude-control requirements.

Requirements for the velocity change thruster:

• Total impulse: 3,000 to 7,000 newton-seconds.
• Thrust: 445 newtons maximum.
• Input power during burn: less than 1 watt.
• Specific impulse: 280 seconds.

Requirements for the attitude control thruster:

• Total impulse: about 2.4 newton-seconds.
• Minimum impulse bit: 0.044 newton-seconds.
• Response time: less than 0.005 seconds.
• Pulse rate: 1 Hz.

It turns out that the velocity change and attitude control thrusters can have independent systems. We propose a new innovation whereby the nanosat launch vehicle propulsion system also serves double duty as the velocity change thruster system, and this can be done without having to carry the propulsion energy source into orbit. This can only be achieved via laser propulsion in which the laser beam energy that is used to launch a nanosat into orbit is also used to provide velocity change thrust in orbit. This novel innovation dramatically reduces the mass, size, cost, and complexity of nanosats because they will only need to carry minimal onboard attitude control thrusters and propellant to carry out routine, minor attitude adjustments. The innovative nanosat laser propulsion concept is presented in Chapter 2.

Miniaturized solid propellant gas generators could be used as attitude control thrusters. Forty-eight 50 millinewton-second pulses are required to reorient the nanosat after it achieves the required orbital altitude. Although this could be achieved either by a monopropellant or a cold gas thruster, it could also be achieved using an array of gas generators. Such miniaturized gas generators have already been successfully built and commercialized by companies such as MOOG and Lockheed-Martin Space Systems. By incorporating micro-electromechanical systems techniques, the devices have been produced relatively inexpensively. Miniaturized electric propulsion attitude control thrusters, such as pulsed plasma and MEMS field-emission electric propulsion thrusters, have been developed and are now emerging into widespread commercialization.

(The remaining sections of Chapter 1 — guidance, navigation and control; command and data handling; power systems; thermal; RF communications; mechanical and structures; instruments; ground systems; and autonomy — set out the mass, power and radiation budgets for each nanosat subsystem, and are omitted for length; the complete text is at the source.)

Chapter 2: Laser Lightcraft Nanosatellite Propulsion

Laser propulsion is a new and exceptional method for reaching space. By launching spacecraft on a beam of electromagnetic radiation, researchers will have developed the first new method of achieving orbit since the late 1950s. In this concept, a remote or ground-based energy source, such as a ground- or space-based laser beam generator, transmits power to a spacecraft via a beam of electromagnetic radiation. The spacecraft collects the beam energy and uses it to power the propulsion system. This concept has the advantage of using the ambient air as the working fluid in the atmosphere and carrying propellant only for use outside the atmosphere, leaving the energy source for heating the propellant on the ground. This results in a tremendous weight reduction and improved performance benefit for the spacecraft because a large propellant mass and heavy energy source are not carried onboard.

The laser-propelled vehicle, called "Lightcraft" because it flies on a beam of laser light, is designed to harness the energy of a laser beam and convert it into propulsive thrust. In the earliest laser-propelled rocket designs, beamed energy from a ground-based laser, with near-visible wavelengths, is absorbed by a heat exchanger onboard a rocket, and is transferred to a working fluid. The heated fluid — hydrogen, ammonia and the like — then produces thrust by expansion through a nozzle as in a conventional chemical rocket. An alternative to this scheme is to use the beamed energy to ablate an onboard solid propellant, such as Delrin, to generate thrust. However, a more recent incarnation of this concept, developed by the Air Force Research Laboratory (AFRL) at Edwards AFB, California, is for the Lightcraft to operate in two propulsion modes: airbreathing, by detonation wave, and rocket ablation, by deflagration. The Lightcraft operates in airbreathing mode up to Mach 5 and 30 km altitude, and in laser thermal rocket mode, using liquid, gaseous, or Delrin ablation propellant, in space. Figure 1 shows the Air Force X-25LR, 25 cm diameter, Lightcraft concept. The Air Force X-50LR Lightcraft has twice the diameter of the X-25LR.

In the two-mode propulsion concept, a forebody aeroshell acts as an external compression surface for the airbreathing engine inlet. Affixed to the bottom of the craft is a parabolic-shaped afterbody mirror, which serves as a primary receptive optic for the laser beam and as an external plug nozzle expansion surface. The primary thrust structure is the centrally located annular shroud, which provides air through the inlet and also acts as a ring-shaped energy absorption and propulsion chamber for plasma formation. The air inlet is closed when the Lightcraft operates in the rocket mode.

The Lightcraft is very lightweight and uses its shape to facilitate vertical flight. The craft has the appearance of a fat acorn when viewed from the side. The lower portion of the craft is a very highly polished metal mirror, whereby the lower point of the acorn shape is the midpoint of a stretched-out parabolic mirror. The Lightcraft receives kilojoule pulses from a ground-based infrared laser at a rate of 25 times per second. The axisymmetric, off-axis parabolic collection mirror facilitates flight by concentrating the pulsed laser light into an annular focus. The laser beam's pulse interacts with the mirror, spreading out and focusing into an annular area inside the circumference of the craft. The intensity of the 18 microsecond pulsed laser is sufficiently high that atmospheric breakdown occurs in the annular area, causing inlet air to momentarily burst into a highly luminous plasma at 10,000 to 30,000 K, thereby producing a superheated plasma shock wave, with instantaneous pressures reaching tens of atmospheres, that generates thrust in the direction of the laser beam. A lip around the craft's circumference, akin to a plug nozzle, directs the expansion of the plasma, creating downward thrust expansion. Multiple laser pulses and an atmospheric refresh of breakdown air generate the flight. This airbreathing pulsed-detonation engine concept owes its origins to the German V-1 buzz bomb of the Second World War, which ran on aviation fuel.

For the purpose of this report, we envision a Lightcraft Earth-to-Orbit (ETO) transportation system that operates according to the following scenario. The airbreathing engine mode develops quasi-steady thrust by pulsing at a variable rate that depends on the Mach number and altitude flown along the flight trajectory to orbit. Once the Lightcraft reaches very high altitude and climbs above the atmosphere, it begins to operate in the thermal rocket mode using onboard propellant to convert and expand the laser energy for propulsion. The Lightcraft is spin-stabilized and can be launched vertically upward or on a slant upward trajectory, hover in mid-air, and undergo powered descent and landing. The ground-based laser beam generator system consists of the following: a power supply; a high-power, megawatt-class laser beam generator and transmitter using novel beam optics; and automated tracking, hand-off and safety systems.

History of the Lightcraft Technology Demonstration Program

The laser Lightcraft project originally grew out of the Lightcraft Technology Demonstration Program funded by the Strategic Defense Initiative Organization (SDIO) Laser Propulsion Program in the late 1980s. In the 1990s, a joint program involving the NASA Marshall Space Flight Center and the Propulsion Sciences and Advanced Concepts Division of the AFRL Propulsion Directorate developed and tested an experiment to determine the feasibility of using high-power pulsed lasers to launch a spacecraft into orbit. Successful tests at the White Sands Missile Range High Energy Laser Systems Test Facility demonstrated the first passively controlled vertical free flight of an object that was propelled by the U.S. Army's 10 kW Pulsed Laser Vulnerability Test System infrared CO₂ laser. Laser boost capability was demonstrated at the facility with a Lightcraft reaching 43 m vertically in 2-second gyroscopically stabilized free flights, which was followed by horizontal guide-wire flights of 121.9 m lasting 10 to 20 seconds. A subsequent series of test flights achieved an altitude of 38.7 m. L. Myrabo, in a private communication from Rensselaer Polytechnic Institute, Troy, New York, 2009, recently reported vertical Lightcraft test flights achieving 68 m altitude.

This achievement can be compared to the first successful flights of Robert Goddard's liquid propellant chemical rocket, which attained a height of 12.5 m after a 2.5 second burn in March 1926. In sharp contrast with Goddard's rockets, there is absolutely no fuel on board the prototype Lightcraft, which has a diameter of 10 cm, mass of 20 to 40 g, and is machined from a solid block of 6061-T6 aluminum. Five different Lightcraft designs have been flight-tested using the pointing and tracking system on the Pulsed Laser Vulnerability Test System laser. Current Lightcraft designs are limited to about 60 g mass and 15 cm in diameter by that laser. A megawatt-class laser will be necessary for a larger kilo-class Lightcraft to reach orbit, and components for these lasers exist, which would demonstrate the feasibility of this technology for low cost access to space.

Figure 2. AFRL Test Vehicle in Vertical Flight. Figure 3. Time-Lapse Photo of a Lightcraft Undergoing an Outdoor Vertical Flight Test. Figure 4. Lightcraft Flight-Test Vehicle Used in Horizontal Guide-Wire Flight Tests — the top of the vehicle is to the right and the laser beam strikes the stretched-out parabolic mirror and propulsion section on the left. Figures 5 and 6. Lightcraft Undergoing Horizontal Guide-Wire Flight Test.

Summary of Technical Performance and Benefits

We outline below the propulsion performance features of the laser Lightcraft launch system:

• The system is single-stage-to-orbit and completely reusable.
• Almost no onboard propellant is required, the reaction mass being free air, except for the small internal amount of propellant needed for final ascent to orbit and orbital maneuvering.
• Vehicle specific impulse is essentially infinite, being several thousand seconds or more in rocket mode.
• Payload mass fractions are 50 to 95 per cent.
• These systems are simple, reliable, safe, environmentally clean, and could have a very high all-azimuth on-demand launch rate.
• They reduce space launch costs by two to three orders of magnitude below today's levels: estimated launch costs are 20 dollars per kg to 600 dollars per kg of payload, not including life cycle and launch operations costs.
• The feasibility and physics principles have been proven by the AFRL's Lightcraft concept Demonstration Program.

Lightcraft systems have sufficient power density to operate as Earth-to-orbit launch systems. It requires a beam power of 0.1 to 1 MW per kg of vehicle mass, while orbit-to-orbit propulsion requires a modest 0.1 to 10 MW of total beam power. The ground-based megawatt-class laser beam generator is state-of-the-art technology. The cost of generating electrical power for the ground-based laser beam generator is about 10 cents per kilowatt-hour, which translates to under 2 dollars per kg of payload. An SDIO study showed that all launch to orbit conditions for a Lightcraft could be satisfied by a single, high-power ground-based laser — with or without the aid of a low altitude laser relay mirror or space-based laser beam generator system. The majority of the system mass required to launch a payload to orbit is left on the ground in the form of the beam generators and their electrical power sources. The dry spacecraft mass can be further reduced by two orders of magnitude, and thus the operating costs reduced by a factor of 10, to under 2 dollars per kg of payload, if buckytubes are used to construct the vehicle and its subsystems.

Lightcraft Nanosatellite Configuration

The Lightcraft nanosat configuration consists of: a conically shaped forebody for lift and aerodynamic compression of ingested airflow, prior to its detonation by laser heating during atmospheric flight; an annular cowl or shroud within which air detonation or propellant ablation, by intense laser heating, occurs; and a parabola-shaped afterbody whose mirrored surface focuses beamed laser energy into regions of sufficient smallness for intense air or propellant heating to occur. The vehicle is powered by laser airbreathing propulsion, by detonation of air, until hypersonic speed within the sensible atmosphere is reached; and then the vehicle is powered by laser rocket propulsion, by heating of propellant, during flight above the sensible atmosphere, until cut-off velocity for orbital flight is reached.

The low vehicle propellant fraction for laser powered Lightcraft, about 0.5 of vehicle takeoff mass, resulted in vehicle takeoff masses that were approximately 45, 80, and 360 times less than those of conventional rockets for placing masses of 10 kg, 5.0 kg, and 1.0 kg into low Earth orbit (LEO). And preliminary life-cycle cost estimates made during the AFRL study by Froning and Davis indicated that transportation system costs for placing 10 kg, 5.0 kg, and 1.0 kg of mass into orbit using Lightcraft and ground-based lasers would be approximately 3, 5, and 15 times less than with conventional rockets.

One of the two most important findings from the Froning and Davis study is the significant influence of Lightcraft drag on airbreathing laser propulsion performance, and the consequence of this on laser rocket propulsion performance during the latter phase of Lightcraft flight. A significant reduction in both Lightcraft size and drag coefficient, as compared to that of the initial government baseline design, was needed for acceptable airbreathing thrusting acceleration during atmospheric flight. Both size and drag coefficient reduction were accomplished in several steps — with both size and drag reduction accomplished during the first step, and further drag reduction, by increased forebody fineness ratio, during the second step.

It was also found that sufficient Lightcraft airbreathing thrust required thrust variation with altitude, somewhat comparable to that achievable by contemporary airbreathing propulsion systems, whose flight dynamic pressure and thrust remain constant with increasing vehicle altitude and speed until constant dynamic pressure can no longer be maintained. Here, acceptable airbreathing thrust minus drag performance was needed to reach maximum airbreathing speed, Mach 10, within acceptably short flight times and distances. And such short times and distances were required to ensure adequate receipt of beamed power by the Lightcraft out to the longest ranges associated with laser rocket propulsion flight, where beamed power would travel the longest distances through the atmosphere and space, and collected power would drop to lowest values.

Froning and Davis also determined that the ground-based laser selected — wavelength 1.62 micrometres, 10 MW radiated power, 10 m diameter aperture — would enable a Lightcraft takeoff mass of 8 kg and Lightcraft propellant mass of 4 kg. Therefore, this would allow approximately 4 kg of mass to be placed into orbit with the selected ground-based laser. And vehicle synthesis work determined that the remaining masses for the Lightcraft airframe, propulsion, and control systems would be 0.63 kg, 0.46 kg, and 0.45 kg, respectively, together with a 30 per cent contingency of 0.47 kg.

Froning and Davis further indicated that small commercial off-the-shelf chemical propulsion systems, with sufficient thrust, would be about a factor of 7 to 12 heavier than those needed to meet Lightcraft orbit circularization needs. However, such mass reductions were deemed possible with emerging MEMS technologies being developed under the National Nanotechnology Initiative for both chemical and field-emission electric propulsion thrusters. It was also found that the currently configured composite structure for the Lightcraft forebody must be reduced from 2-ply to 3-ply, with the same ply-thickness, to meet Lightcraft airframe mass requirements.

Another important finding in the study was the significant influence of the ground-based laser wavelength on Lightcraft performance. The beam propagation geometry associated with Earth-to-orbit laser propulsion by means of ground-based lasers is adverse. Beam propagation distances through the Earth's atmosphere are short during initial flight phases when the path length traveled by laser energy to the Lightcraft is least. But during latter flight phases, when the vehicle itself is above the sensible atmosphere, the beam propagation path within the atmosphere is much longer, and power losses due to atmospheric attenuation become ever greater with increasing range. And since power losses due to laser beam spreading — even in vacuo — also increase with increasing distance from the laser, power losses are greatest at the end of laser propulsion, when vehicle distance from the laser is greatest.

For a ground-based laser with given aperture diameter, adaptive optics, atmospheric conditions, and radiated power, the laser power collected by the Lightcraft was found to be extremely sensitive to laser wavelength. Wavelength determined the amount of radiated laser power lost through thermal blooming, turbulence, and extinction during beam passage through the Earth's atmosphere, in addition to the power lost from diffraction, that is beam spreading at longer ranges, during propagation through the vacuum of space. And since each loss mechanism was a function of wavelength, Froning and Davis considered each loss mechanism in their estimation of lost power for the six different laser wavelengths associated with the six different ground-based laser candidates that were evaluated in the study.

A significant fraction of laser-radiated power is lost at maximum laser propulsion range, when the necessary cut-off velocity for orbital flight is achieved, even if there were no atmospheric transmission losses at all. And additional losses associated with beam propagation through the atmosphere result in power losses on the order of 75 per cent to 99 per cent. Significantly more power would be available at the end of laser airbreathing flight than at the end of laser rocket flight. This might benefit surface-to-air Lightcraft missions that would mainly entail airbreathing flight.

Figure 12. Captured Laser Power versus Increasing Range from an 11.2 micrometre CO₂ Laser. For a given laser aperture diameter, adaptive optics, and atmospheric conditions, the collected power decreases with increasing range, shown for a vertical laser-pointing angle and for a final laser-pointing angle of 83 degrees from the vertical that occurs at maximum laser propulsion range, about 500 km, where the Lightcraft reaches maximum speed.

Figure 13. Influence of Lightcraft Range and Pointing Angles on Captured Power. The comparison is between the selected laser wavelength of 1.62 micrometres and the 11.2 micrometre CO₂ laser wavelength chosen for a government baseline Lightcraft, for the highest radiated power, 10 MW, and the largest laser aperture, 10 m, that was deemed practical for Air Force operations and systems, with a Lightcraft capture diameter of 30 cm.

Unfortunately the demonstrated laser beam power levels for the attractive 1.62 micrometre wavelength, which suffered the least propagation losses, are relatively modest. This attractive laser wavelength is associated with the wavelength-tunable free-electron laser (FEL), whose maximum beam power is currently in the 20 kW range. Thus, there is the need for a 500-fold increase in FEL beam power to achieve the 10 MW beam power required for 10 kg-class Lightcraft Earth-to-orbit propulsion. However, 100 kW beam FEL designs are being proposed by the Navy for prototyping and testing in FY 2011 and 2012.

The physics and technology of FELs will allow beam power to be scaled up to 1 MW or higher as long as thermal loading of the beam optics and electron losses in the electron beam recirculation loop can be mitigated using engineering solutions. Beam combining of several 1 MW or higher FELs can achieve a total combined beam output power of 10 MW or higher. Other newly emerging high-power laser technology that shows promise for achieving megawatt-class beam power includes bulk slab solid-state and high-power fiber lasers; the former has already achieved over 100 kW of beam power while the latter is getting close to it. Present megawatt-class lasers that are based on available proven technology include a proposal for a five-beam, 2.5 MW per beam, electron gun-driven CO₂ and gas mixture laser which combines five laser beams to achieve 10 MW of total beam output power. These systems are described further in Chapter 4.

Life Cycle of Lightcraft System

Froning and Davis found that ground-based laser costs comprised the major portion of a Lightcraft Earth-to-orbit transportation system — with ground-based laser costs comprising about 80 per cent of the total laser Lightcraft system life-cycle cost. The life-cycle cost of a laser Lightcraft Earth-to-orbit transportation system was estimated using Lightcraft vehicle and ground-based laser cost inputs from AFRL together with programmatic cost inputs from another cost database. Table 1 shows the programmatic assumptions together with the system acquisition and operation costs for the various Lightcraft vehicle and ground-based laser system elements. Laser acquisition and operation costs were assumed to be shared with another user and all operations costs are reduced to one-half those values estimated from historical data. Launch costs are seen to be extremely low, only 74,141 dollars per flight, with laser-associated costs comprising approximately 92 per cent of the laser-powered Lightcraft Earth-to-orbit transportation system life-cycle cost.

(Table 1, the Laser Lightcraft Model Cost Summary, assumes a 10-year mission model at 1,000 launches per year with a 2.0 kg payload per launch, and gives a total program cost of 741.4 million dollars, of which 680.4 million is development and acquisition — 624.8 million of that being the 10 MW ground-based laser — with an average cost per kilogram, based on operations costs, of 3,052 dollars. The table is omitted for length; the complete text is at the source.)

Chapter 3: Laser Lightcraft Weapon Mission Selection Study

The objective of the Froning and Davis study was to examine crucial future Air Force launch vehicle missions and select at least one that might be performed very cost-effectively by Lightcraft vehicles powered by beamed electromagnetic energy from airborne or ground- or sea-based lasers. It entailed identification and analysis of promising launch vehicle missions for laser-propelled Lightcraft and assessments of the identified and analyzed missions by experts in the mission areas.

It was concluded that the most promising Air Force mission for a laser-propelled Lightcraft is the placement of Earth and space observing nanosats of up to 3 kg mass into LEO. Such a laser-propelled Lightcraft would also serve as a "Lightsat", because it would use the Lightcraft's laser propulsion optics as a telescope for observing military targets on Earth and in space. Additional estimated mass for performing the Lightsat function is no more than about 1 kg, if the Lightcraft forebody structure panels can be unfurled in orbit and used as solar power collectors. Such a Lightcraft system appears capable of reaching LEO at one fifth to one tenth the cost required for placing a similar Lightsat system into LEO using multistage chemical rocket systems.

Figure 14. Ground- and Sea-to-Space Concept: appropriate rotation of a high-energy laser beam, emanating from a ground- or sea-based laser, guides and propels an integral Lightcraft pico- or nano-satellite along an Earth-to-orbit ascent path until orbital conditions are reached, or until beamed laser energy can no longer be transformed into Lightcraft thrust.

Assumed Laser and Lightcraft Limitations

Beamed power levels achievable with envisioned high-energy laser technology are assumed to be no more than about 10 MW for ground- or sea-based lasers, and no more than about 2.0 MW for the much lighter and smaller airborne lasers installable on Boeing 747, B-1 Lancer, or C-130 subsonic aircraft. Lightcraft takeoff masses no more than about 20 kg can be accelerated to orbital velocities by maximum ground- or sea-based laser power levels, and Lightcraft takeoff masses no more than about 4 kg to 8 kg, depending on the magnitude of velocity and acceleration needed, can be accelerated to very high velocities by the lower allowable masses and power levels of airborne lasers.

Lightcraft Trajectory and Mission Limitations

• Small allowable angle between centerlines of the laser beam and Lightcraft vehicle axes before significant thrust reduction occurs.
• Small allowable angles of attack, that is the angle between Lightcraft centerline and velocity vector, before significant thrust reduction occurs.
• Limited capability for engaging multiple missile or aircraft threats in an allowable time interval because of relatively long laser beam-riding time against each threat.

Additional Lightcraft hardware needed for some tactical missions includes terminal seeker guidance for hit-to-kill accuracy against missiles and maneuvering aircraft; axial and lateral propulsion and control for missile and maneuvering aircraft interception; and additional mass along the Lightcraft centerline for hardened target penetration.

Lightcraft Mission Investigations: Summary and Conclusions

If laser propulsion can provide nearly all the velocity change a Lightcraft needs to reach LEO, then Lightcraft nanosat systems, which combine both launch vehicle and nanosat subsystems within a single vehicle, may be achievable with launch masses in the 2 kg to 10 kg range; and such Lightcraft nanosat systems appear capable of reaching LEO at one fifth to one tenth the cost required using multistage chemical rocket systems.

If hypersonic magnetohydrodynamic airbreathing propulsion research and development currently underway at the National Science Foundation, NASA and the AFRL is successful, then Lightcraft dry masses as heavy as 100 kg can be launched from aircraft flying at Mach 10 to 12 at about 30 km above the Earth. Such Lightcraft could be propelled by laser power as high as 100 MW that can be generated from the electrical power of ionized-air-slowing by interacting electric and magnetic fields within hypersonic magnetohydrodynamic airbreathing engines.

Sufficient impact energy for destruction of high-speed ballistic missiles above the atmosphere is possible with chemical propulsion and uncooled infrared detectors, for semi-active homing and axial and lateral acceleration during end-game, integrated into Lightcraft vehicles for an approximate 100 per cent dry mass increase, from 1.0 kg to 2.0 kg. But multiple target interception within allowable time is limited by the relatively long beam-riding time needed for the Lightcraft to reach and destroy each target.

Although laser-propelled Lightcraft appear capable of performing certain Air Force tactical missions, and are much less expensive than missiles currently used for such missions, the laser and aircraft costs associated with Lightcraft launches are much greater. Also, clouds impair Lightcraft air-to-ground and air-to-air effectiveness, while air-to-air and air-to-space effectiveness is limited by long Lightcraft beam-riding times. Thus, no truly attractive Lightcraft combat mission was found. On the other hand, Lightcraft were found to be extremely attractive, compared to chemical rockets, in boosting microsats, nanosats and picosats to LEO, whereby the Lightcraft plus ground- or sea-based laser costs are significantly less than multistage chemical rocket costs. Thus, the selected Lightcraft missions are launch vehicle missions involving ground, sea and air launches of Lightcraft to LEO, with air launch occurring at either subsonic or hypersonic speed.

Lightcraft Ground- and Sea-to-Space Investigation

The current Air Force Lightcraft vehicle concept has been designed for not only placing nanosats into LEO at low cost, but also for performing much of the nanosat function as well. In this concept the precision optics system that focuses ground- or sea-based laser light into the Lightcraft's cowl area for propulsion is also used as a space telescope for viewing military targets on Earth and in space. And structural panels on the Lightcraft forebody are also used as solar panels that are unfurled in orbit for generation of satellite power. Thus, the current Lightcraft's design allocates only 0.1 kg of its 1.0 kg dry mass for exclusively nanosat functions. There is a military need for 1.0 kg to 2.0 kg nanosats with optical sensors for visual inspection of unknown objects in space and on Earth.

Since conventional expendable rockets could conceivably be an alternative to laser-powered Lightcraft for the rapid placement of military nanosats in LEO, a cursory comparison of Lightcraft and conventional rockets was made by Froning and Davis to get some idea of their comparative costs. Hybrid rocket sizing and costing was based upon tactical strategic missile sizing and costing information. Lightcraft sizing assumed a propellant mass fraction of 0.5 and 1.0 MW of laser power per pound of payload, that is dry mass, placed into orbit. Costs for laser power and refurbishment were based upon AFRL estimates, amortized over a fewer number of flights. Although these Lightcraft and laser costs are higher, being based upon much fewer flights, than those of previous AFRL estimates, they are believed to be consistent with the conventional rocket costs, and therefore applicable for relative cost comparisons. More detailed future Lightcraft and conventional rocket designs and cost comparisons are, of course, needed before a strong argument can be made for either design.

The estimated costs, assuming 100 flights over a 10-year period, indicate that Lightcraft could boost nanosats in the 2.0 kg to 5.0 kg range into LEO at about one tenth to one fifth the cost of expendable rockets. But Lightcraft cost superiority over conventional rockets is less overwhelming for satellites that are significantly heavier. Thus Lightcraft appear extremely attractive for satellite delivery missions only if Lightcraft dry masses, including the satellites being carried, are less than about 5.0 kg.

(Tables 2 and 3, giving stage-by-stage masses, specific impulses and propellant fractions for three-stage hybrid rocket launch vehicles alongside the single-stage Lightcraft — whose effective specific impulse is 1,452 seconds at a propellant fraction of 0.5, for an impulsive velocity of 9.84 km/s — are omitted for length; the complete text is at the source.)

Future Nano- and Pico-Satellite Mission Concepts

Coherently cooperating swarms are a novel innovation for replacing structures with information by placing many formation-flown small satellites into a loose swarm and causing them to cooperate coherently. This is very different from the so-called distributed small satellite LEO constellations currently pursued, in which individual small spacecraft perform essentially the same functions as larger satellites but at lower spacecraft mass and cost.

In contrast, the swarms cooperate coherently and form a real distributed system in which the whole is more than the sum of the parts. A generic description would be a constellation of small spacecraft each performing its separate function, but these functions combine to create at a central location a much larger virtual spacecraft, or sensor aperture, that exists solely because of the cooperation of the spacecraft. Each of a large number of small satellites in a swarm or other constellation will radiate or receive signals and combine them in phase, or coherently, regardless of their actual location in orbit. This creates coherent RF or optical apertures that are essentially unlimited in size.

Each satellite's position is only crudely station kept, and the satellites adjust the time delay or phase delay of the signals they repeat to compensate for their position errors, causing their repeated signals to add coherently at a collection point. This technique can be easily applied to RF transmitters and receivers, and with more demanding accuracy to optical transmitters and receivers. The result in either case is a large swarm or loose constellation of satellites that act as one large antenna or optical array, even though they are separate and their positions are neither constant nor lie along a parabola or plane in space. The individual satellites can be as simple as one-element flying chips or as complex as today's self-contained sensing spacecraft of various sizes.

The advantage of coherently cooperating distributed systems is that they can form sparse RF antennas and optical sensors with diameters so large that they would be impossible to implement with filled apertures even if formed with adaptive membranes, and have orders of magnitude smaller mass. The relative locations of individual spacecraft in the swarm can be controlled by MEMS field-emission electric propulsion, tethers, or by cleverly conceived orbits in which the elements of the array appear to orbit a common center within it, thus eliminating the need to use propulsion at least for first order station keeping.

These array functions can be made coherent over very great distances. RF antennas with sizes of hundreds of kilometers and optical telescopes of hundreds of meters diameter can be formed. These systems can enable new capabilities not possible with single spacecraft either acting alone, as a proliferated but non-coherent constellation, or as relays for each other. Formed of nanosats and picosats, such swarms will contain so many spacecraft that the economics of true mass production will come into play in space for the first time, greatly reducing the cost of producing the system. In addition, these systems feature the advantages of truly distributed satellite systems, including fault tolerance, robustness, survivability, reconfigurability by software, and the ability to be incrementally emplaced and upgraded as budgets are available.

These swarms can be implemented in a cost effective manner using laser propulsion for both launch and orbital insertion. However, the system designs are flexible enough to allow for the use of alternative conventional launch vehicle technologies. The technologies to produce these swarms and their constituent nanosats or picosats probably can be demonstrated by 2015 and deployed in space by 2020.

Rotating Picosat Swarm Array Radio Frequency Collector

An unconventional, large sparse antenna array RF collector spacecraft with a small surface footprint even when deployed in geosynchronous Earth orbit (GEO) separates different sources in proximity and also detects weak signals. Its implementation would result in a highly desirable, long dwell RF emitter detection capability.

At the heart of this system is a large antenna that is formed by a swarm of tiny elements that make up the lens of a space-fed array with no structure. The antenna is a sparse, self-cohering array formed from a large number of picosats rotating, in relative coordinates, in a plane around a central orbital point in GEO. The picosats are self-contained repeater spacecraft. Each one receives the ground signal, delays it, and retransmits the signal so that it arrives at the feeds at the same time as a direct ray through the center of the array. The time delay of each picosat is self computed based on its location in the swarm, as measured by a local differential global positioning system navigation signal, to compensate for its deviation from its assigned ideal location. Each picosat digitizes, delays, frequency shifts, and retransmits its received signals independently, causing an in-phase composite signal from the ground to be received at the feeds.

The relative positions of these picosat elements change slowly, and only small and infrequent propulsive maneuvers are needed for constellation maintenance. A tether along the local vertical at the central point holds the receivers and navigation reference at the focus against a counterweight. A pseudorandom distribution of the picosats suppresses the antenna grating lobes, and intensive computation greatly reduces much of the remaining sidelobes, creates multiple beams, and steers the ensemble of the individual beams anywhere on Earth. The antenna system will function with far fewer elements as a more sparse array, though with limited sensitivity. This system can be incrementally emplaced, upgraded, and even funded with capability growing as budget is available, as opposed to the usual all-or-nothing functioning of today's spacecraft.

The antenna size is 20 km by 40 km and contains 150,000 picosats, each of which weighs 23 grams. The feed array is held in position by a 50 km long, lightweight tether against a counterweight. There is no truss or other structure. Each picosat is gravity-gradient stable, has a dipole array facing Earth, and a broader beam antenna array facing the receivers.

The effective collecting aperture of the array is equal to that of an equivalent 80 m diameter filled aperture antenna. The coverage spot diameter can be varied by choosing the diameter of the array that is active, with spot sizes on Earth as small as 30 m at 10 GHz, 300 m at 1 GHz, or 3 km at 100 MHz. It can receive sub-watt signals from individual cell phones. The entire constellation weighs 3,500 kg, but that could be reduced in the future to 35 kg if buckytubes are used to construct the system.

High Resolution Surface Sampling Radiometry

Highly sensitive radiometry at low microwave frequencies with a small ground footprint would result in high resolution microwave radiometry sampling maps of soil moisture and other surface characteristics, as well as passively detected larger targets. The constellation and array implementation follows that of the preceding concept, except that it is designed to map the surface radiation rather than detect discrete emitters. The antenna size is 8 km by 12 km and contains 12,000 picosats, each of which weighs 23 grams. The feed array is held in position by a 40 km long, lightweight tether against a counterweight. The constellation scans its coverage spot electronically in a 1,200 km zig-zag swath from its 4,000 km orbit by modulating the time or frequency shift of the ensemble of picosats.

The effective collecting aperture of the array is the sum of those of the picosats, and in this example, equal to that of an equivalent 11 m diameter antenna. However, the coverage spot diameter is set by the total aperture diameter of 8 km by 12 km, and thus is 100 m at 1 GHz. Five constellations would result in a 5 hour global revisit with zig-zag coverage of the scanned swaths. The entire constellation weighs 3,000 kg, but that could be reduced in the future to 30 kg if buckytubes are used to construct the system.

High Resolution Surface Mapping Radiometry

Highly sensitive radiometry at low microwave frequencies with a very small footprint on the ground would result in high resolution microwave radiometry maps of soil moisture and other surface characteristics and passively detected larger targets, with 100 per cent of Earth's surface mapped with a 5 hour revisit time. The principle of operation is the same as that of the previous concept, except that a multiple element detector array is used in a pushbroom scanning mode for complete Earth coverage rather than only sampling coverage.

The implementation is similar to that of the preceding concept, except that tethers hold a receiving array that must be 2 km long to obtain the 1,200 km instantaneous swath width with a resolution of 100 m. It consists of a 2 km long focal surface with 12,000 printed dipoles, shaped into a focal surface by gravity gradient forces balanced against magnetic forces from a superconducting conductor around its periphery, acting on a piezoelectric, electron beam-shaped, adaptive membrane substrate. The large antenna is formed by a swarm of tiny elements making up the lens of a space-fed array.

The antenna is a 4 km by 6 km diameter, sparse, self-cohering array formed from 12,000 picosats weighing 23 grams each, rotating in relative coordinates in a plane around a central orbit point. Their locations are initially selected to lie in a plane, and their spacings are pseudorandom to minimize the sidelobe levels, with each picosat designed to loosely stationkeep inside a box 10 m on a side. The relative positions of these picosat elements change slowly, and only small and infrequent stationkeeping propulsive maneuvers are needed for constellation maintenance.

The effective collecting aperture of the array is the sum of those of the picosats, and in this concept, equal to that of an equivalent 6 m diameter antenna at 2 GHz. However, the coverage spot diameter is set by the total aperture diameter of 4 km by 6 km, and thus is 100 m at 2 GHz from a 4,000 km orbit. Five constellations would produce 100 per cent global coverage with 5 hour revisit for time critical measurements. The entire constellation weighs 3,000 kg, but that could be reduced in the future to 30 kg if buckytubes are used to construct the system.

Rotating Nanosat Swarm Distributed Radar

An extremely powerful space-based radar, this concept would allow detection of most air, land, sea, and space targets, as well as many low observable targets anywhere, with one or a few constellations in GEO. A large, sparse array antenna using a swarm of nanosats creates a space-based radar system. The implementation is similar to that of the preceding rotating swarm concepts, except that it generates and radiates extremely large peak and average powers, and given the generally high angles of viewing can detect and track many air, space, and surface targets from GEO.

The constellation is composed of 10,000 nanosats that are self-contained repeater spacecraft weighing about 1 kg each. Each nanosat receives the ground signal, digitizes, delays, and retransmits it, causing it to arrive at the feeds at the same time as a direct ray through the center of the array. The time delay of each nanosat is self computed based on its location in the swarm, as measured by a local navigation signal, to compensate for its deviation from its assigned ideal location. Commands for beam sweep delays are superimposed on the time delays of each nanosat. Each nanosat generates 10 W of average power and 10 kW peak power at 0.001 duty cycle, has a helical film antenna that increases its gain and doubles as a solar sail for infrequent stationkeeping maneuvers, and has a tether for coarse gravity gradient stabilization. A 50 km tether supports the feed, transmitter, and navigation reference assembly against a counterweight. The antenna lens is 2 km by 4 km.

The effective area of the array is the same as that of a 50 m diameter filled aperture. The total effective RF radiated power of the system is 3 GW peak and 3 MW average. Although specific performance calculations have not been done for this concept, these powers are so large that the radar should have the sensitivity from its location in GEO for detecting and tracking many targets simultaneously, and most low observable targets as well, because they are all designed and oriented so as to have their low observables in near-horizontal directions. Three constellations would provide essentially complete global coverage. The entire constellation weighs about 11,000 kg in GEO — this could be reduced in the future to 110 kg if buckytubes are used to construct the system — and can be emplaced and replaced incrementally using laser-powered Lightcraft launch vehicles or even small conventional launch vehicles. It could even be funded incrementally.

Simple, Distributed, Hyperspectral Sensor

This concept presents an unconventional method of implementing a hyperspectral sensor of great spectral and spatial resolution. Its implementation would allow the detection of very many spectral intervals simultaneously, and it has a small field of view from GEO so that the instrument can dwell on and resolve particular targets of interest. It also has a large field of regard so that one spacecraft covers a significant fraction of a hemisphere.

The concept uses a Fresnel zone plate, which is oriented roughly parallel to the local horizontal just below GEO. It is supported by a tether that extends well above the GEO altitude, and may or may not have a counterweight at the top end. The gravity gradient causes the ensemble to remain Earth pointing along the local vertical, with its center of mass in GEO.

The Fresnel zone plate has a long focal length, and thus the surface and ring locations can be imprecise compared with conventional optics. In addition, the lens is a thin film membrane and will be light and inexpensive. It is highly frequency dispersive, and thus its focal length is a sensitive function of wavelength. Small, self-contained optical sensor nanosats are placed on the tether at many locations, with each nanosat's optics filtered for response at only that narrow spectral region focused at its distance from the lens. The nanosats can transmit directly to the ground or their signals can be combined in one transceiver, also on the tether.

This system has a 100 km long tether, which weighs only a few kilograms in GEO. The Fresnel zone plate is 100 m in diameter, has a collecting aperture equivalent to a 30 m filled aperture, and requires only a surface accuracy of centimeters in the visible light region. It is constructed of thin film with deposited aluminum rings and is an adaptive piezoelectric membrane kept flat by an electron beam in response to an optical figure sensor. MEMS field-emission electric propulsion thrusters are at the sensor's periphery for attitude control, with 1,000 nanosats attached to the tether. Each nanosat is a self-contained optical sensor spacecraft, integrated into a 0.1 kg package. The sensor thus detects 1,000 wavelengths simultaneously. The sensor has a resolution as small as 40 cm on the ground from GEO. Its field of view can be scanned over an area by tilting the ensemble using the thrusters on the nanosats, or libration modes can be excited in the tether so that the field of view scans across a 2,000 km coverage area in a quasi-random mode, eventually covering the entire area. These thrusters also serve for stationkeeping. Its total weight is less than 900 kg in orbit, which can be reduced in the future to 9 kg if buckytubes are used to construct the system.

(The ballistic missile defence analysis of the air-to-space Lightcraft, and the recommended Lightcraft missions section that follows it, are omitted for length; the complete text is at the source.)

Chapter 4: Summary of Multi-Megawatt Laser Study for Lightcraft Propulsion Applications

Summary of Laser Study Performed by Textron Systems

In 2002, V. Hasson of Textron Systems Corporation conducted a study of candidate multi-megawatt laser systems for laser Lightcraft propulsion. Candidate high-power and high-energy lasers identified in the study:

• Carbon dioxide (CO₂) laser: technical issues include large wavelength and atmospheric absorption of the laser beam.
• Carbon monoxide (CO) laser: technical issues include large wavelength, atmospheric absorption of the laser beam, and toxicity of the lasing fuel.
• Hydrogen fluoride or deuterium fluoride laser: technical issues include atmospheric absorption of the laser beam, corrosive lasing fuel chemicals, pulse energy, running cost, and beam quality.
• Chemical oxygen-iodine laser: technical issues include fuel chemicals, pulse energy, and running costs.
• Bulk slab solid-state laser: technical issues include cost, average power, and run duration.

The first four gas dynamic and chemical laser candidates have already demonstrated megawatt-class average beam output power. The study then reviewed the development, testing and operational legacy of the first candidate laser technology, which included a review of the various system architectures that use other gas mixtures combined with CO₂. The study recommended a new design for a 10 MW beam output power electron gun-driven CO₂ and gas mixture laser because this technology does not require additional research and development and can be implemented now. The other gases selected for the lasing fuel are nitrogen and hydrogen, which, in combination with CO₂, offer superior performance over systems using helium.

However, very recent technologically disruptive innovations led to a tenfold increase in the beam output power of bulk slab solid-state lasers and their newly emergent solid-state cousin, called high-power fiber lasers, which has made these devices more competitive with high-power chemical and gas dynamic lasers on the basis of average beam power, peak beam power, electrical wall plug and optical efficiencies, cost, complexity, mass, and size. Free-electron lasers are another class of laser technology that was not reviewed by Hasson, but recent technological innovations are accelerating their development to the point where their present average beam output power of 20 kW will be increased to 100 kW or higher within the next 12 to 24 months following the publication of this report. The output beams of several of these laser devices can be optically combined to produce a single beam with megawatt-class average output power.

An outline of the conceptual design features of the proposed 10 MW electron gun-driven CO₂ and gas mixture laser is:

• Scalability of total beam output power, beam combining concept.
• Power oscillator or master oscillator-power amplifier design.
• Unstable optical resonator cavity with grating and rotating mirrors beam-combine techniques.
• Flow and gas handling system with blow down and exhaust to the atmosphere.
• Acoustics suppression with expansion horn downstream and anode muffler.

In this concept there are four separate laser transmitters each generating 2.5 MW output beams that are combined into a single 10 MW output beam. The oscillator parameters for each beam transmitter are: energy loading of 300 joules, with higher loadings at reduced gas temperature; gain volume 0.27 cubic metres for each of four lasers; gain length 3 m; specific laser output 65 joules per litre; estimated extraction efficiency 20 per cent; pulse repetition rate 125 Hz at 20 microseconds; laser power 2.5 MW per beam across four beams for 10 MW; laser energy per pulse 18 kJ per beam across four beams for 72 kJ; output wavelengths 10.6, 10.2, 9.6 and 9.3 micrometres mixed; gas mixture ratio for nitrogen, carbon dioxide and hydrogen of 3 to 1 to 0.08; and gas pressure of one atmosphere.

(The remainder of Chapter 4 — the optical resonator and gas flow system specifications, the survey of emergent high-power solid-state, fiber and free-electron laser technologies, and the high energy laser beam control section covering pointer-trackers, shared aperture systems and adaptive optics — is omitted for length; the complete text is at the source.)

Chapter 5: Conclusion

To reduce mission costs, advanced technology components and a novel laser propulsion system can make nanosats, and picosats, compact, lightweight, low power, and low cost. By producing a large quantity of nanosats for a given mission, the per-unit cost will be reduced to a small fraction of satellite procurements for traditional missions. Mission operation costs will be minimized by the incorporation of both onboard and ground autonomy, use of heuristic systems, and use of a novel laser propulsion system to launch the nanosats into LEO. Laser propulsion is an enabling technology in which a laser-propelled vehicle, called a Lightcraft, harnesses the energy of a high-energy laser beam and converts it into propulsive thrust.

The laser-propelled Lightcraft is an Earth-to-orbit transportation system that develops quasi-steady airbreathing thrust by pulsing at a variable rate along the flight trajectory to orbit, and then when it climbs above the atmosphere it begins to operate in the thermal rocket mode using onboard propellant to convert and expand the laser energy for propulsion. The Lightcraft is spin-stabilized and can be launched vertically upward or on a slant upward trajectory, hover in mid-air, and undergo powered descent and landing. The system is single-stage-to-orbit and completely reusable with no onboard propellant required, the reaction mass being free air, except for the small internal amount of propellant needed for final ascent to orbit and orbital maneuvering. MEMS field-emission electric propulsion thrusters could provide onboard attitude and stationkeeping propulsion. The Lightcraft specific impulse is essentially infinite, several thousand seconds in rocket mode, while payload mass fractions are 50 per cent to 95 per cent.

Laser-propelled Lightcraft systems are simple, reliable, safe, environmentally clean, and could have a very high all-azimuth on-demand launch rate. This novel propulsion system reduces space launch costs by two to three orders of magnitude below today's levels, with estimated launch costs of 20 dollars per kg to 600 dollars per kg of payload, not including life-cycle and recurring launch operations costs. The entire Lightcraft launch system is comprised of a ground, sea, or airborne laser beam generator consisting of a power supply, a high-power megawatt-class laser beam generator and transmitter using novel beam optics, and automated tracking, hand-off and safety systems.

The most promising military mission for laser-propelled Lightcraft is the placement of Earth and space observing nano- and pico-satellites of up to 3 kg mass into LEO. Such Lightcraft could also serve as a Lightsat, because it would use the Lightcraft's laser propulsion optics as a telescope for observing military targets on Earth and in space. Such a Lightcraft system appears capable of reaching LEO at one fifth to one tenth the cost required for placing a similar Lightsat system into LEO using multistage chemical rocket systems. Other potential missions include using laser-propelled Lightcraft as ground-, sea- or airborne-launched kinetic kill weapons to shoot down enemy ballistic missiles. Very innovative near-term missions could also include deploying Lightcraft nano- and pico-satellites to form swarms of small spacecraft which cooperate coherently to form a real distributed system in which the whole is more than the sum of the parts. This would be a constellation of small spacecraft each performing its separate function, but these functions combine to create at a central location a much larger virtual spacecraft, or sensor aperture, that exists solely because of the cooperation of the spacecraft. This creates coherent RF or optical apertures that are essentially unlimited in size, which could offer unprecedented high-resolution radiometry, hyperspectral imaging, radar, and RF interception.

Launching laser-propelled Lightcraft nano- and pico-satellites to LEO requires megawatt-class lasers. Textron Systems Corporation's proposed 10 MW electron gun-driven CO₂ and gas mixture laser is a multi-megawatt-class system that can be implemented now because this technology requires little or no additional research and development. This system offers realistic near-term, low-cost Lightcraft launch capability. However, this system is large, requires a large amount of gas propellant to fuel the laser, and the system infrastructure will cost over 200 million dollars.

The newly emergent bulk slab solid-state, high-power fiber, and free-electron laser technologies being explored by the various Department of Defense directed energy weapons programs offer higher electrical-to-optical efficiencies and overall laser performance, compact and portable system size, less complexity and smaller weight, all at much lower system and infrastructure cost. These lasers are scalable to megawatt-class beam power, and so we roughly estimate that the overall system and infrastructure cost to deploy such laser systems to launch a Lightcraft to LEO will be from several factors to an order of magnitude, or more, lower than for the electron gun-driven CO₂ and gas mixture laser system.

Removing the waste heat produced by high-power laser systems is an important factor driving the physical limitations of scaling up the beam output power. An innovative matched-refractive-index liquid is used to rapidly remove the heat produced by a 150 kW bulk slab solid-state laser weapon, while the very high surface area-to-volume ratio of high-power fiber lasers allows for the rapid removal of heat from the gain medium without the need for external cooling. Phase-change materials are being explored, and devices using such materials have recently demonstrated the ability to store very large quantities of the waste heat produced by high-power solid-state lasers, which is a different way of rapidly removing large amounts of heat from the solid-state gain medium. Unlike solid-state laser systems, free-electron lasers are not affected by heat problems, while their gain medium, a vacuum, cannot be damaged.

Launching a laser-propelled Lightcraft nanosat or picosat from the ground, sea, or air into LEO requires controlling and steering the high-energy laser beam, while at the same time making real-time adjustments to account for platform motion, optical train and atmospheric effects on beam propagation, so that the beam maintains high quality, low-loss, precision contact with the Lightcraft during the entire flight. Recent technical innovations in optical train design and other system architecture have evolved beam control devices for high-energy laser weapons toward new implementations. New beam control devices and high-power optical train combinations have a resulting beam line that is considerably simpler, smaller and lighter than current architectures. Almost every component in the beam line performs multiple functions, thereby dramatically reducing the high-power optical component count. This approach also packages all beam control sensors, processors and drivers into a single turret assembly.

In 2005, the AFRL Propulsion Directorate at Edwards AFB, California, concluded their laser Lightcraft propulsion research, development, test and evaluation program before launching a Lightcraft test vehicle into LEO was demonstrated. At present, the Air Force Office of Scientific Research is funding the Brazilian Air Force's hypersonic shock tunnel study of laser Lightcraft propulsion in collaboration with Leik Myrabo's laser propulsion group at Rensselaer Polytechnic Institute, Troy, New York. This author recommends that the Department of Defense, in collaboration with NASA, return laser Lightcraft propulsion research and development to the United States and restart the space launch test flight demonstration program of the Air Force X-50LR Lightcraft, which was originally proposed by Dr. Frank Mead.

(The thirty-five numbered references are omitted for length; the complete text is at the source.)