AAWSAP DIRD, Laser Lightcraft Nanosatellites, November 2010

UNCLASSIFIED//fOR. 8FFl&IAk W&li ,n11av 01 November 2010 ICOD 30 August 2010 DIA-08-1011-001 Defense Intelligence Reference Document Defense Futures Laser Lightcraft Nanosatellites Laser Lightcraft Nanosatellites UNCLASSIFIED,C/F8R 8FFl@lslct tl!II! 8HL I UNCLASSIFIED//P'8R 8FFl01iflk W&li QtU.¥ The Defense Intelligence Reference Document provides non-substantive but authoritative reference information related to intelligence topics or methodologies. Prepared by: Technology Warning Division (DW0-4) Defense Warning Office Directorate for Analysis Defense Intelligence Agency 611thnr· AAP Person 58 Administrative Notes: (U) COPYRIGHT WARNING: Further dissemination of the photographs in this publication is not authorized. This product is one in a series of advanced technology reports produced in FY 2010 under the Defense Intelligence Agency, Defense Warning Office's Advanced Aerospace Weapon System Applications.---.... (AAWSA) Program. Comments or questions pertaining to this document should be addressed to I -·--. ) MP Person AAP Perso~ ______ I AAWSA Program Manager, Defense Intelligence Agency, ATTN: JUIAF - DI/DWO-3, 1 ·-··­ 1 ---· Bldg 6000, Washington, DC 20340-5100. ii UNCLASSIFIED//liQQ OFFIQIA.. W&E &•U:Y UNCLASSIFIED//f8R 8FFIGI/Jk W&li QIIL¥ Contents Chapter 1: Nanosatellite Technologies .................................................................. 3 Chapter 2: Laser Lightcraft Nanosatellite Propulsion ..............................................11 Chapter 3: Laser Lightcraft Weapon Mission Selection Study ..................................27 Chapter 4: Summary of Multi-Megawatt Laser Study for Lightcraft Propulsion Applications .....................................................................................................42 Chapter 5: Conclusion ......................................................................................68 References .......................................................................................................71 Figures Figure 1. Air Force X-25LR Laser Lightcraft ..........................................................12 Figure 2. AFRL Test Vehicle in Vertical Flight ........................................................14 Figure 3. Time-Lapse Photo of a Lightcraft Undergoing an Outdoor Vertical Flight Test 15 Figure 4. Lightcraft Flight-Test Vehicle Used in Horizontal Guide-Wire Flight Tests ...... 16 Figure 5. Lightcraft Undergoing Horizontal Guide-Wire Flight Test ...........................16 Figure 6. Lightcraft Undergoing Horizontal Guide-Wire Flight Test ...........................17 Figure 7. Lightcraft Concept .......... ...... ................ ... .......... ........... .......... ............ .18 Figure 8. Lightcraft Trajectory and Associated Pointing Angles .................................19 Figure 9. Lightcraft Vehicle Evolution ... ................. ..... ........ ............. .......... ...........20 Figure 10. Attenuation Effects on Captured Laser Beam Power................................22 Figure 11. Influence of Trajectory and Laser Wavelength on Captured Power .............22 Figure 12. Captured Laser Power vs. Increasing Range from 11.2 µm CO2 Laser ........23 Figure 13. Influence of Lightcraft Range and Pointing Angles on Captured Power........24 Figure 14. Ground/Sea-to-Space Concept .............................................................27 Figure 15. Air-to-Space Concept . .................................................. ............ ......... .28 Figure 16. Schematic of Power Oscillator Optics ....................................................44 Figure 17. Schematic of MOPA ............................................................................45 Figure 18. Schematic of the Laser N2/CO2/H2 Gas Flow System ................................45 Figure 19. Northrop Grumman's Joint High Power Bulk Slab Solid-State Laser ...........48 Figure 20. DARPA's High Energy Liquid Laser Area Defense System .........................50 Figure 21. Phase Change Materials Allow Storage of Large Intermittent Heat Loads While Slow Regeneration Removes Heat from Aircraft ............................................52 Figure 22. Typical HPFL MOPA Design... ...............................................................54 Figure 23. Fiber Laser Beam Combining Techniques ...............................................54 Figure 24. Pumping Fiber Lasers .........................................................................56 Figure 25. Large and Small Diameter Fiber Lasers .................................................56 Figure 26. Single Mode Fiber Laser Modules ..........................................................57 Figure 27. Multimode HPFLs ................................ ...............................................57 Figure 28. Free-Electron Laser............................................................................58 Figure 29. Free-Electron Laser Mechanism ............................................................58 Figure 30. Free-Electron Laser Electron Beam Phase-Space Evolution .......................59 Figure 31. Recirculating-Beam FEL System ...........................................................60 Figure 32. High-Power FEL Optical Resonator ........................................................62 Figure 33. Notional Long Range HEL Beam Control System .....................................64 UNCLASSIFIED,tfFQA QFFIGIAI:: 1!181: 8HLY iii UNCLASSIFIED/ /fOR. 8FFl&IAk W&li ,HIL:¥ Figure 34. HEL Beam Pointer/Tracker...................................................................65 Figure 35. Basic Shared Aperture Beam Control System .........................................66 Figure 36. HEL Adaptive Optics System................. ...... .........................................67 Tables Table 1. Laser Lightcraft Model Cost Summary ......................................................25 Table 3. Estimated Costs for Hybrid Rocket and Lightcraft Launch Vehicles for ETO Table 4. Influence of Target Velocity and Intercept Angle on Impact Energy and Table 5. Influence of Lightcraft and Target Velocity on Impact Energy and Required Table 2. Performance and Estimated Weights for a Hybrid Rocket and Lightcraft ........30 Flight..............................................................................................................32 Required Mass ..................... ...................... .. .................. ...................................33 Mass ............................................................................................. ..................35 UNCLASSIFIED'Fl'FAR OFFJCJ0L: Pl&li ,>all.¥ iv UNCLASSIFIED//fOR. 8FFl&IAk W&li ,>all.¥ Laser Lightcraft Nanosatellites 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 (MASINT) or surveillance and eavesdropping (SIGINT) 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 $500k. By producing a large quantity of nanosats for a given mission, the per-unit cost will be reduced to a small fraction of UNCLASSIFIED//FOR 061ilCl.li.k W&& 8Ptl'l 1 UNCLASSIFIED/ /P'OR. 8FFIEiIAk W&i QPd~¥ 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. UNCLASSIFIED,'fFQA QFFIEiIAL 1!181!! OHLY 2 UNCLASSIFIED//fOR. 8FFl&IAk W&li ,HIia¥ 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 (C&DH) 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: 1) each nanosat must raise its orbit apogee to the appropriate radius, 2) and 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 (1W) and attitude-control (ACS) requirements. Requirements for the Av Thruster: • Total impulse: 3,000 to 7,000 N-sec. • Thrust: 445 N maximum. • Input power (during burn): < 1 watt. • Specific impulse: 280 seconds. Requirements for the ACS Thruster: • Total impulse: $ 2.4 N-sec. • Minimum impulse bit: 0.044 N-sec. • Response time: < 0.005 sec. • Pulse rate: 1 Hz. It turns out that the tiv and ACS thrusters can have independent systems. We propose a new innovation whereby the nanosat launch vehicle propulsion system also serves double duty as the tiv 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 tiv 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 ACS thrusters and propellant to carry out routine, minor attitude adjustments. The innovative nanosat laser propulsion concept is presented in Chapter 2. UNCLASSIFIED//FOR O61ilCl.li.k W&& 8Ptl'l 3 UNCLASSIFIED//fOR. 8FFl&IAk W&li ,HIia¥ Miniaturized solid propellant gas generators could be used as ACS thrusters. Forty­ eight 50 mN-sec 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 (MEMS) techniques, the devices have been produced relatively inexpensively. Miniaturized electric propulsion ACS thrusters, such as pulsed plasma and MEMS field-emission electric propulsion (MEMS FEEP) thrusters, have been developed and are now emerging into widespread commercialization. GUIDANCE, NAVIGATION AND CONTROL Guidance Navigation and Control (GN&C) subsystem key technologies and concepts have been identified to enable successful altitude determination of spin-stabilized and three-axis-stabilized nanosats for future missions. They include miniaturization of a sun sensor and horizon crossing indicator. The miniature precision "fan" sun sensor will pinpoint the sun virtually everywhere in the entire celestial sphere with every satellite rotation. The sun sensor will be required to weigh less than 0.25 kg, draw less than 0.1 watt, operate on no greater than a 3.3 volt bus, and meet a 0.1° resolution requirement. The miniature horizon crossing indicator has a small bore-sight field of view that is mounted at an angle off the spin axis. As the spacecraft rotates, a cone of coverage is formed. The sensor must be capable of detecting Earth over a range of orbital radii with a pointing accuracy of 0.05°. Total horizon crossing indicator weight and power will be less than 0.2 kg and 0.1 watt, respectively. Of particular interest to Constellation missions is the incorporation of GPS onboard the nanosats, to eliminate ground-based ephemeris generation. This allows for increased autonomy and simpler, more accurate time resolution onboard the spacecraft. For GPS to fit within the constraints of a nanosat, the receiver electronics need to be miniaturized into a layer within the C&DH module. COMMAND AND DATA HANDLING Developing the C&DH subsystem for a nanosat presents some unique challenges, with low mass (0.25 kg) and low power (0.5 W) requirements being the biggest drivers. Advanced microelectronic solutions are being developed to meet these challenges. The microelectronics developed must be modular and of scalable packaging to both reduce cost and meet the requirements of various missions. This development will utilize the most cost effective approach, whether infusing commercially driven semiconductor devices into spacecraft applications or partnering with industry in the design and development of high capacity data processing devices. The major technologies will include: lightweight, low power electronics packaging; radiation hard, low power processing platforms; high capacity, low power memory systems; and radiation hard, reconfigurable, field programmable gate arrays (RHrFPGA). The C&DH requirements are as follows: • Power: 0.5 watts. UNCLASSIFIED,<,<FOA QFFI&IAL HSI!! OHLY 4 UNCLASSIFIED//fOR. 8FFI61Ak U&F QJsll.¥ • Weight: 0.25 kg. • Input data rate: 2 kbits/sec. • Output data rate: 100 kbits/sec. • Data storage: 2 Gbits. • Encoding: advanced convolutional. • Processing speed: 12 MIPS. • Radiation tolerance: > 100 krads total dose. In order to develop a low mass C&DH, a lightweight and low power electronics packaging method must be used. The packaging method that will be chosen must have a small volume and small footprint (6 cm x 6 cm x variable height). The packaging technique must provide data on programmable substrates and data on a compliant interconnects for space use. A multi-chip module (MCM) has been successfully produced by Pico Systems Inc. A combined effort to reduce mass, power, size and cost led to the development of the CMOS Ultra Low Power Radiation Tolerant (CULPRiT) system on a chip, and "C&DH in your Palm" are technologies that enable the power reduction required for nanosats. The goals of these technologies are a 20: 1 power reduction over current 5-volt technology, foundry independence of die production, and radiation tolerance. Another technology enabling a decrease in volume is the RHrFPGA, which reduces volume by replacing many logic functions/circuits with one die. The RHrFPGA also allows concurrent design by decoupling the logic design from the module, shortens the design schedule, lowers the part count, and eases rework. The above technologies allow for higher levels of electronic integration, effectively combining spacecraft subsystem electronics and instrument electronics into the smallest possible mass, power, and volume. POWER SYSTEMS Total spacecraft power is limited by the small satellite size. The Sun's power density is 1.35 kW/m 2. Assuming 15% conversion efficiency for a 0.3 m x 0.1 m disk shaped nanosat (cross section of 0.03 m2), with a 67% area coverage, this results in a total electric power of only 4.0 watts. Lightweight, efficient solar array panels that minimize the effective array mounting area are needed. Dual or triple junction GaAs solar cells that give 18% conversion efficiency at end of life (EOL), and assuming a more optimistic area factor of 85%, will result in only 6.2 Wat EOL. Small satellites that do not have extended solar panels simply do not intercept a large solar power density and must use the available power very efficiently. For a small spinning satellite, it is expected that three solar cells will be connected in series along the spin axis, and groups of three will be connected in parallel around the circumference. Each section will generate 3.3 volts and rotate into and out of sunlight as a unit. Voltage drops at 3.3 volts, bus regulation, circuit protection (e.g., fuse or circuit breaker) and Lithium ion battery discharge characteristics are being studied. Highly elliptical orbits in the ecliptic plane where the apogee velocity is very low will cause a several hour eclipse during part of the year. Spacecraft batteries to cover this eclipse period presents a significant mass impact. However, only a 10° orbit plane UNCLASSIFIED//FOR 061ilCl.li.k U&& 8Ptl'l 5 UNCLASSIFIED//fOR. 8FFI61Ak U&i 0'911.¥ inclination relative to the ecliptic, will reduce the maximum eclipse period to about one hour. Inclusion of spacecraft batteries is then justified. Passive thermal control will be used to keep the spacecraft electronics within 10°C of ambient temperature, and hence will not require electric power for heating. Using such a scenario, a battery requirement of about 2 amp-hours at 3.3 volts will allow full spacecraft functionality during an eclipse. Twelve AA size Lithium-ion batteries meet the requirement and only weigh 480 grams. Circuits that have high current demands, such as thruster solenoids and fuses, need to be augmented with components that have a lower power density than batteries, but also have lower internal resistance. Ultra-capacitors are being explored for this application. Miniaturization of the power system electronics (PSE) to meet the weight and size requirements of the nanosats is a considerable challenge. The ideal approach is to eliminate the PSE completely, by having a fixed electrical load and batteries provide the needed bus regulation. This yields a simplified system consisting of solar cells, batteries, and minimal circuitry. A more immediate approach to miniaturization is to produce hybrid modules that measure approximately 5.08 cm x 3.17 cm x 1.27 cm and weigh 100 grams for each PSE component, namely the solar array regulator, battery regulator, and low voltage power converter. The combination of these three components into one module will reduce the size and weight another order of magnitude. THERMAL Although an inclination change by 10° renders maximum shadows below two hours, we evaluate the case of a maximum eight hour shadow for the purpose of generality. Three thermal configurations are considered: (1) top and bottom of the nanosat are insulated, the inside of the cylindrical solar array is not insulated, allowing internal heat transfer between the internal equipment and the array; (2) the entire nanosat is insulated, top and bottom as well as inside the solar arrays, except for a radiator on top, sized to radiate the internal electrical dissipation; and (3) the internal equipment is thermally isolated as well as possible from an "outside shell" with a controllable two­ phase heat transport device which can be "shut off" during Earth shadows, serving as the only thermal coupling between the equipment and a radiator on the outside surface. The key advantage of configuration (1) is its reliability, or robustness. Since the temperature of the nanosat is set by a high energy balance (heat in - heat out) dominated by the absorbed solar energy, the operational temperature of the nanosat is relatively insensitive to top and bottom multilayer insulation (MLI) properties, or, largely, to internal heat dissipation. However, the feature that yields the operational reliability, i.e., the high energy balance, also results in a rapid drop in temperature when the solar load disappears during the Earth shadow. During the maximum eight hour eclipse used for this evaluation, it was found that internal temperatures dropped by about 60°C, which would result in internal temperatures in the range of - 30°C to - 40°C. At the same time, the solar arrays dropped to a temperature of about 60°C. Based on past experience, these end-of-eclipse temperatures are reasonable. UNCLASSIFIED,<,<FOA QFFIEiIAL HSI!! OHLY 6 UNCLASSIFIED//fOR. 8FFI61Ak U&F QJsll.¥ Because configuration (2) has a much smaller overall energy balance than configuration (1), it is much more sensitive to MLI properties and to internal power dissipation. However, eclipse performance improves. During the ~ 8 hour eclipse, internal temperatures drop by only 20°C, a marked improvement, with end-of-eclipse temperatures well within the range of most spacecraft components. It should be noted that the solar arrays, since they are now isolated from the body of the nanosat, drop to temperatures of about - 110°C. Even these solar array temperatures should not pose a problem. For example, the solar arrays of many geosynchronous satellites drop routinely to temperatures of about -150°C during the 72 minute eclipse experienced by these spacecraft at each equinox season. The key feature of configuration (3) is that the equipment is coupled to an external radiator only with a two-phase heat transport device, such as a capillary pumped loop (CPL) or loop heat pipe (LHP). Operational temperatures are again maintained to temperatures of about 20°c nominal with a properly sized radiator. However, the temperature is also totally dependent on the proper operation of the two-phase "loop." The two-phase heat transport device can be made redundant by the addition of a second loop if single fault tolerance is desired. Note that redundancy is not a consideration for the other two configurations. During the ~ 8 hour eclipse, further improvement is realized, with internal temperatures dropping by as little as 6°C if the internal payload is well insulated from the exterior of the nanosat. As in configuration (2), the solar array temperatures drop to about -110°C. For certain equipment or science instruments, the temperature control afforded by this type of "active" design may be necessary. A moderate amount of technology development has been underway since 2000 to enable a two-phase heat transport system for use in a nanosat. The small size and low heat transport requirements of the nanosat will necessitate significant downsizing of today's flight qualified two-phase systems. This reduction will be accomplished by leveraging recent successful tests of a small cryogenic two-phase CPL. RF COMMUNICATIONS The onboard RF subsystem must be small, low mass, and low power. The system specifications are: • Mass: 0.5 kg. • Power consumption: 0.5 watt. • Transmission data rate: up to 100 kbits/sec. • Command reception data rate: 1 kbit/sec. • Range: 3 to 5 Earth radii. • Channel type: BPSK. • Effective isotropic radiated power: 0.15 watt (-8.2 dbW). • Carrier frequency: 8,470 MHz. The tracking system should be coupled with this communication subsystem to maximize efficiency in mass and power. UNCLASSIFIED//FOR O61ilCl.li.k U&& 8Ptl'l 7 UNCLASSIFIED//fOR. 8FFI61Ak U&i 0'911.¥ The communications subsystem is further complicated by constellations requiring spin­ stabilized nanosats. A spinning nanosat cannot easily point an antenna toward Earth. Therefore, a low gain omni antenna is assumed and communications must take place near perigee, when the range is 3 to 5 Earth radii. A large ground antenna and high data rate compression must be used to achieve reasonable data rates with minimum power. This places an additional burden on the ground stations for both sensitive receivers/bit synchronizers and advanced decoders. These same considerations limit data rate for satellite-to-satellite communication. Although the inclusion of an onboard command receiver is highly desired, it puts an additional strain on an already challenged nanosat mass and power budget. For this reason, the concept of a totally autonomous, receiverless nanosat design appears most attractive. However, "receiver-on-a-chip" technology has advanced to the point where including a receiver onboard looks feasible. The biggest disadvantage of a receiver now becomes the ground personnel and software needed to support the ability to command the nanosat. Command actions taken onboard will of course be limited to basic functions such as "transmit data" because of the lack of redundancy and mechanical functions. Although scenarios have been defined to allow nanosats to autonomously determine when to transmit their stored data, utilizing a receiver to control the telemetry downlink from the ground still has value. The capability of uploading flight software changes, as well as sending a master reset if necessary, would also exist with such an onboard command receiver. MECHANICAL AND STRUCTURES The nanosat mechanical system will be kept as simple as possible. The ideal nanosat mechanical design should consist of a one-piece structure on which all other components are mounted. Multifunctional structures can provide thermal control, shielding and serve as substrates for printed circuit boards. For example, diamond facesheet honeycomb panels can serve as a structure, thermal conductor and radiator, and printed circuit board substrates. The diamond facesheet provides ten times greater thermal conductivity than aluminum and can dissipate heat from high power density electronics modules with a low mass comparable to carbon fiber composites. Another example is the structural battery system. It consists of a honeycomb panel whose core is filled with the cells of a nickel-hydrogen battery (or other flight qualified cell technology). Concurrent engineering and fabrication techniques will be used to create a single computer model for the design, analysis (structural, thermal, and dynamic), and fabrication of the nanosat and its components. Dynamic modeling capabilities to simulate nanosat deployments will provide faster designs and a reduction in the amount of deployment testing required. This approach will significantly lower development costs by reducing duplication of effort, chances of errors, the number of drawings and paperwork required. Mass production techniques not traditionally used for spaceflight hardware will be used, such as casting and injection molding. Options being considered for the nanosat structure material are: cast aluminum; cast aluminum-beryllium alloy; injection molded UNCLASSIFIED,<,<FOA QFFIEiIAL HSI!! OHLY 8 UNCLASSIFIED//fOR. 8FFl&IAk W&li ,HIia¥ plastic; fiber reinforced plastic; flat stock composite construction; and carbon nanotubes (a.k.a. "Buckytubes") or carbon nanotubes composited with other materials. The material will be selected based on mass, cost, manufacturability, ease of assembly and integration, and suitability for the space environment. Streamlined testing is needed for up to 100 or 1000 nanosats per mission. Performing a complete test program on each unit would be prohibitively expensive and time consuming. We need to reduce the quantity of testing required while assuring product quality to meet program cost and schedule goals. Lot testing and statistical quality control methods should be developed to verify quality and structural performance by testing a small subset of the total number of nanosats. INSTRUMENTS Instruments for in-situ and remote measurements must be miniaturized to fit within the mass and volume constraints of a nanosat. Power consumption must also be scaled down accordingly. Instrument sensitivities cannot be compromised in the process. Instrument electronics need to be combined with nanosat subsystem electronics to achieve higher degrees of integration yielding reduced mass and volume. Instrument software will be designed to evaluate the onboard data and adjust instrument data rates and modes to efficiently capture the data of highest priority. GROUND SYSTEMS The large number of nanosats in a constellation is a challenge to the ground system in getting all of the data to the users. In a typical baseline mission, there are times when up to ten (or more) nanosats would be within communications range of a ground station at a single time. A minimal model for the ground station contacts shows that they can support a nanosat constellation with only two ground stations located on opposite sides of the Earth. The schedulers will prioritize the contacts, with the nanosats in the higher period orbits getting priority. Nanosats in the lower period orbits have more opportunities to dump their data, and therefore can have lower priority without risking any data loss. Since the nanosats are autonomous, the operations concept for a mission requires only a few operators to determine the nanosat orbits, schedule the ground stations, and to investigate anomalies on the spacecraft. Automated systems will monitor the housekeeping data from the spacecraft and they will flag problems for the spacecraft engineers to investigate. The large number of nanosats allows the risk management to be different for this mission than for single spacecraft missions. Except for commands to initiate the data downlink, the ground system will not command the nanosats for normal operations. The only commands that the ground system sends would be program loads to resolve or work around problems and failures. The large number of nanosats in a constellation is a configuration control challenge for the data tracking, the schedules, the command loads, the science or intelligence data, and the engineering data. The ground system will use IDs, colorcoded user interfaces, and other techniques to ensure that the operators and users can keep track of the data UNCLASSIFIED//FOR 061ilCl.li.k W&& 8Ptl'l 9 UNCLASSIFIED//fOR. 8FFl&IAk W&li ,>all.¥ associated with a particular nanosat. Constellations that fly in close formation can benefit by the use of inter-nanosat communications to reduce ground station contention. The data would flow from a single nanosat to the ground instead of coming from every nanosat. Communications protocols for inter-nanosat communications must be developed. AUTONOMY Support costs are high if single-satellite mission operations and data analysis practices are scaled to a constellation mission. Autonomy onboard the spacecraft and on the ground is therefore required to ensure that mission objectives are efficiently and inexpensively met. Nanosat autonomy will make use of onboard and ground-based remote agents with the overarching goal of maximizing the scientific or intelligence return from each nanosat during the mission lifetime. The remote agents achieve this goal by monitoring and appropriately controlling nanosat subsystems. Additionally, the onboard agent monitors the full complement of spacecraft sensors and instruments to heuristically separate scientific or intelligence events of interest from background events, thereby intelligently fitting the science/intelligence data within allocated spacecraft storage resources. Nanosats with distant orbits are out of communications range of a ground station for nearly a week. Nanosat subsystems could be compromised if faults occurring during this blackout period were not readily addressed. An unacceptable loss of scientific or intelligence data could also occur. Therefore, the onboard agent will incorporate the capability to detect, diagnose, and recover from faults. Certain failure scenarios may not be correctable by the onboard agent. These faults will be deferred to the ground agent for handling. Each nanosat will include data in its telemetry on the health and status of each subsystem and a history of commands autonomously issued since the last ground contact. The ground system will then attempt to diagnose problems based on this data. Additionally, collective knowledge of actions taken by all nanosats in the constellation will reside within the ground system by virtue of the data dumps made during each contact. From this data the agent can detect trends and systematic conditions not otherwise observable onboard the nanosat. These highly autonomous systems will present a unique set of challenges not only to the system designers, but also to those involved in spacecraft testing. Careful consideration must be given to the design of the test program to ensure that the state­ space of the remote agents is validated and verified. It is equally important to implement this program in a cost-effective manner. However, we could likely justify deploying considerable resources to address this issue since the methods developed to solve these challenges can be applied to numerous missions. UNCLASSIFIED,<,<FOA QFFI&IAL HSI!! OHLY UNCLASSIFIED//fOR. 8FFl&IAk W&li ,>all.¥ Chapter 2: Laser Lightcraft Nanosatellite Propulsion Laser propulsion is a new and exceptional method for reachi ng space. By launching spacecraft on a beam of electromagnetic radiation, researchers will have developed the first new method of achieving orbit since the late 1950's. 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 [ 1-8]. 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, etc.) 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, CA, is for the Lightcraft to operate in two propulsion modes: airbreathing (detonation wave) and rocket ablation (deflagration). The Lightcraft operates in air breathing mode up to Mach 5 and 30 km altitude, and in laser thermal rocket mode (using liquid, gaseous, or Delrin ablation propellant) in space [7, 8, 9-16]. Figure 1 shows the Air Force X-25LR (25 cm diameter) Lightcraft concept. The Air Force X-SOLR Lightcraft has twice the diameter as the X-25LR. UNCLASSIFIED//FOR 061ilCl.li.k W&& 8Ptl'l 11 UNCLASSIFIED//fOR. 8FFl&IAk W&li ,HIia¥ Figure 1. Air Force X-25LR Laser Lightcraft {courtesy of F. Mead, AFRL/ PRSP, Edwards AFB, CA). 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/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 (see Figure 2). 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 (10,000 - 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 UNCLASSIFIED,<,<FOA QFFI&IAL HSI!! OHLY 12 UNCLASSIFIED//fOR. 8FFI61Ak U&F QJsll.¥ 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 WW II 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: 1) power supply; 2) high-power (megawatt-class) laser beam generator/transmitter using novel beam optics; and 3) 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 1980's. In the 1990's, 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 (WSMR) High Energy Laser Systems Test Facility (HELSTF) 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 (PLVTS) infrared CO2 laser. Laser boost capability was demonstrated at the HELSTF 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 (see Figure 2 through Figure 6). A subsequent series of test flights achieved an altitude of 38. 7 m. L. Myrabo (private communication, Rensselaer Polytechnic Inst., Troy, NY, 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 PLVTS laser. Current Lightcraft designs are limited to about 60 g mass and 15 cm in diameter by the PLVTS 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. UNCLASSIFIED//FOR O61ilCl.li.k U&& 8Ptl'l 13 UNCLASSIFIED//fOR. 8FFl&IAk W&li ,HIia¥ Figure 2. AFRL Test Vehicle in Vertical Flight (courtesy of F. Mead, AFRL/PRSP, Edwards AFB, CA). UNCLASSIFIED,<,<FOA QFFI&IAL HSI!! OHLY 14 UNCLASSIFIED//fOR. 8FFl&IAk W&li Qlhll.¥ Figure 3. Time-Lapse Photo of a Lightcraft Undergoing an Outdoor Vertical Flight Test (courtesy of F. Mead, AFRL/PRSP, Edwards AFB, CA). UNCLASSIFIED//FOA QliliHiiilAts W&& 8,.L'l 15 UNCLASSIFIED//fOR. 8FFl&IAk W&li Qlhll.¥ Figure 4. Lightcraft Flight-Test Vehicle Used in Horizontal Guide-Wire Flight Tests (the top of vehicle is to the right and the laser beam strikes the stretched-out parabolic mirror/propulsion section on the left} (courtesy of F. Mead, AFRL/ PRSP, Edwards AFB, CA). Figure 5. Lightcraft Undergoing Horizontal Guide-Wire Flight Test (courtesy of F. Mead, AFRL/PRSP, Edwards AFB, CA). UNCLASSIFIED{fFOA QFFI&IAL ~8! 8HLY 16 UNCLASSIFIED//fOR. 8FFl&IAk W&li ,HIia¥ Figure 6. Lightcraft Undergoing Horizontal Guide-Wire Flight Test (courtesy of F. Mead, AFRL/PRSP, Edwards AFB, CA). 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 is free air), except for the small internal amount of propellant needed for final ascent to orbit and orbital maneuvering. • Vehicle specific impulse (lsp) is essentially infinite (::: several x 103 seconds in rocket mode). • Payload mass fractions are"" 50 - 95%. • These systems are simple, reliable, safe, environmentally clean, and could have a very high all azimuth on-demand launch rate. • Reduces space launch costs by two to three orders of magnitude below today's levels: estimated launch costs are $20/kg to $600/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 [11, 16-25]. Lightcraft systems have sufficient power density to operate as ETO launch systems. It requires a beam power of 0.1 to 1 MW per kg of vehicle mass, while orbit-to-orbit UNCLASSIFIED//FOR O61ilCl.li.k W&& 8Ptl'l 17 --------------- UNCLASSIFIED//fOR. 8FFl&IAk W&li Qlhll.¥ 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 ~ $0.10/kWh, which translates to < $2/kg of payload. An SDIO study [10, 11] 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 < $2/kg of payload), if Buckytubes are used to construct the vehicle and its subsystems. LIGHTCRAFT NANOSATELLITE CONFIGURATION As shown in Figure 7, the Lightcraft nanosat configuration consists of: 1) a conically shaped "forebody" for lift and aerodynamic compression of ingested airflow (prior to its detonation by laser heating during atmospheric flight); 2) an annular "cowl" or "shroud" within which air detonation or propellant ablation (by intense laser heating) occurs; and 3) a parabola-shaped "afterbody" whose mirrored surface focuses beamed laser energy into regions of sufficient smallness for intense air or propellant heating to occur. And as shown in Figure 8, 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. Shroud (Cowl): within which Laser • Laser Airbreathing Flight Heating of Airflow from Zero Velocity to and Propellant Hypersonic Speed Occurs Afterbody: with Mirrored Surface for Focusing Laser Energy into the Laser Shroud (Cowl) \ Beam .______ ___ __ _________ }__ Forebody: for Lift and Compression • Laser Rocket Flight from of Airflow during Hypersonic to Orbital Atmospheric Flight Axi-Symmetric Body Speed Figure 7. Lightcraft Concept 1261. UNCLASSIFIED{fFOA QFFI&IAL ~8! 8HLY 18 UNCLASSIFIED/,'rOR: orrlCIJ!tL USE 014Lf Zenith Ligbtcraft Earth-to-Orbit Maxim um (Cut-Ofl) Velocity Ground Based Laser (Not to Scale) Figure 8. Lightcraft Trajectory and Associated Pointing Angles 1261. The low vehicle propellant fraction for laser powered Lightcraft ( ~ 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 [26] 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. As indicated in Figure 9, a significant reduction in both Lightcraft size and drag coefficient (Co) - as compared to that of the initial government baseline design - was needed for acceptable airbreathing thrusting acceleration during atmospheric flight. Figure 9 shows that both size and drag coefficient reduction were accomplished in several steps - with both size and Co reduction accomplished during the first step, and further Co 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 (q) and thrust remain constant with increasing vehicle altitude and speed until constant q 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. UNCLASSIFIED//FOR 061ilCl.li.k W&& 8Ptl'l 19 UNCLASSIFIED/;'POR: OPPICIAE USE O14Lf Froning and Davis [26] also determined that the ground-based laser selected (). = 1.62 µm, 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 percent contingency (of 0.47 kg). Takeoff Mass =4.0 kg Frontal Area= 0.78 m1 /<3) I TakeoffMass=8.0kg ~ FnmlalArea •.096m' ~ 1.00 i ! ~ = s .10 ... i ... ... ·; ,II...• ~ ~ < '-' l .. 41 lie .. •.. · ...; Q .01 ...·; i.. i ~ 1:1 @ ~ !-­ .001 Takeoff Mass =8.0 kg Frontal Area "' .096 m2 (1) .16~ .0054 (I) Cr,A/W (m21k&) f L .0026 ~ I I I ~ I l I / Range for / Currently ;;iJt' SDIO Llgbtcraft Envisioned &:ramjets - .,, M ... ­ 0 10 20 30 Half-Apell Angle ofForebody (degrees) Figure 9. Lightcraft Vehicle Evolution (in 3 steps) 1261. Froning and Davis [26] further indicated that small COTS 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 UNCLASSIFIED,'fFOA QFFISIAL 1!181!! 8HLY 20 UNCLASSIFIED/;'POR: OPPICIAE USE 014Lf deemed possible with emerging MEMS technologies being developed under the National Nanotechnology Initiative for both chemical and FEEP thrusters (see Chapter 2 for details). 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. Figure 8 illustrates the adverse beam propagation geometry associated with ETO laser propulsion by means of ground-based lasers. It is seen that 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. Here, >.. 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" (beam spreading at longer ranges) during propagation through the vacuum of space. And since each loss mechanism was a function of>.., 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. Shown in Figure 10 (without dimensions) is the fraction of radiated laser power collected by the Lightcraft at maximum laser propulsion range (when necessary "cut­ off" velocity for orbital flight is achieved) for the spectrum of wavelengths investigated. It is seen that a significant fraction of laser-radiated power is lost, even if there were no atmospheric transmission losses at all. And additional losses associated with beam propagation through the atmosphere are seen to result in power losses on the order of 75% to 99%. Figure 11 shows that 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. UNCLASSIFIED'I i'FAA: 061i1Clali.k W&& 8PtLY 21 UNCLASSIFIED//rOR: orrlCIAE USE 014Lf 0 1 :Z 3 4 5 6 7 Laser Waveleogtll/Optimum Laser Wavelength Figure 10. Attenuation Effects on Captured Laser Beam Power [261. 0.4 • Earth-to-Orbit Trajectory 0.3 End of Laser Airbreathing Propulsion o.:z End of Laser Airbreathiog and Rocket 0.1 Propulsion 0 ..__.__ _,_____._.....1..._..____.__ _,__ .._____.__ .....1..._....__.._____.___., 0 2 3 4 6 7 Laser Waveleogtll/Optimum Laser Wavelength Figure 11. Influence of Trajectory and Laser Wavelength on Captured Power [261. 0.5 0.4 0.3 0.2 0.1 0 • Earth-to-Orbit • End ofLuer ~ Alrbreatbiag Trajectory \ and Rocket \ Propulsion \ \ \ Beam Power Attenuation \ due to Difraction \. j Beam Power Attenuation 0... due to Difraction and ', Atmospheric Effects ,....., .............. ............ "o UNCLASSIFIED{fFOA QFFIEiIAL ~81!! 8HLY 22 UNCLASSIFIED/,'rOR: orrlCIJ!tL USE 014Lf Figure 12 shows, for a given laser aperture diameter, adaptive optics, and atmospheric conditions, the decrease in laser power collected by the Lig htcraft with increasing range from a 11.2 µm wavelength CO2 laser. The decrease is shown for a vertical laser­ pointing angle and for a final laser-pointing angle of 83° (from the vertical) that occurs at maximum laser propulsion range (about 500 km), where the Lightcraft reaches maximum speed. 1.E+07---------.-------,-------r------, i- .:= 11.2 micron Laser Wavelength f J:! -:i La er Beam Angle from the Vertical :i 1.E+06-l--½--- ~a:----=-----F""-..::-------:c:-+.,-------::----:---+------t----t 300 400 500 0 100 200 Ligbtcraft Slant Range from Laser (km) Figure 12. Captured Laser Power vs. Increasing Range from 11.2 µm CO2 Laser 1261. Figure 13 shows the significant difference in the laser power collected by the Lightcraft during its laser propulsion phase of flight for the selected laser wavelength of 1.62 µm, and for the 11.2 µm CO2 laser wavelength chosen for a government baseline Lightcraft. This comparison is for a Lightcraft trajectory determined from optimization work during the latter phases of the Froning and Davis study. It was also for the highest radiated power (10 MW) and the largest laser aperture (10 m) that was deemed practical for Air Force operations and systems. Unfortunately the demonstrated laser beam power levels for the attractive 1.62 µm 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 ETO propulsion. However, 100 kW beam FEL designs are being proposed by the Navy for prototyping and testing in FY 2011 and 2012. £ l ! Q,U 1.E+OS-1----- ~ ---,---+-:,--------:----'=l'--..-.::::----f--­ ._ i ... ~ ,:! l .E+OA+----~e-------1-----=---1-----------1 UNCLASSIFIED//FOR 061ilCl.li.k W&& 8Ptl'l 23 6 1s = ..= ~- -= e.o 4 ~ >, ,l:l ] 3 ::I Q. -= u 2 .. 41 t 1 .. GI.. = ~ 0 UNCLASSIFIED/,'rOR: orrlCIJ!tL USE 014Lf •Radiated Laser Power= 10 MW Laser Beam 50• 60° 70° Angle from 75" the Vertical I i I I I IEnd ofLaser1 (Airbreathingl 1.62 ·micron Propulsion : Laser Wavelength I I I I I •Ligbtcraft Lightcraftll I Capture Dia. Achievement of: of30 cm Orbital Speed I i I 11.2 micron : .04 Laser Wavelength \ I MW I __j 100 200 300 400 soo Ligbtcraft Slant Range from Laser (km) Figure 13. Influence of Lightcraft Range and Pointing Angles on Captured Power C261. 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 show promise for achieving megawatt-class beam power include 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 5-beam, 2.5 MW per beam, electron gun-driven CO2/gas mixture laser which combines five laser beams to achieve 10 MW of total beam output power. These systems will be described further in Chapter 4. LIFE CYCLE OF LIGHTCRAFT SYSTEM Froning and Davis [26] found that ground-based laser costs comprised the major portion of a Lightcraft ETO transportation system - with ground-based laser costs comprising about 80% of the total laser Lightcraft system life-cycle cost (LCC). The LCC of a laser Lightcraft ETO transportation system was estimated using Lightcraft vehicle and ground-based laser cost inputs from AFRL/PRSP 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 per flight) with laser- UNCLASSIFIED{fFOA 9FFH31AL ~8! 8HLY 24 UNCLASSIFIED/;'POR: OPPICIAE USE 014Lf associated costs comprising approximately 92% of the laser-powered Lightcraft ETO transportation system LCC. Table 1. Laser Lightcraft Model Cost Summary [261. Laser Lightcraft Model Cost Summary Share Solid-State Laser Cut Ops Costs by 50% Mission Model Length (Years) Launch Rate Per Year Payload Per Launch (kg) Mission F