AAWSAP DIRD Laser Lightcraft Nanosatellites November 1 2010
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Defense
Intelligence
Reference
Document
Defense Futures
01 November 2010
ICOD 30 August 2010
DIA-08-1011-001
Laser Lightcraft
Nanosatellites
Laser Lightcraft Nanosatellites
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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·
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Administrative Notes:
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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 ·-··-
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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
Appl ications .....................................................................................................42
Chapter 5: Conclusion ......................................................................................68
References .......................................................................................................71
Figures
Figure 1. Air Force X-25LR Laser Lightcraft .......................................................... 12
Figure 2. AFRL Test Veh icle 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 Po inting 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. Typ ical 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
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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 2. Performance and Estimated Weights for a Hybrid Rocket and Lightcraft ........ 30
Table 3. Estimated Costs for Hybrid Rocket and Lightcraft Launch Vehicles for ETO
Flight..............................................................................................................32
Table 4. Influence of Target Velocity and Intercept Angle on Impact Energy and
Required Mass ..................... ...................... .. .................. ................................... 33
Table 5. Influence of Lightcraft and Target Velocity on Impact Energy and Required
Mass ............................................................................................. .................. 35
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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. Picosatel lite 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
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satellite procurements for traditional missions. Mission operation costs will be
minim ized by the incorporation of both onboard and ground autonomy and use of
heuristic systems.
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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 traditiona l 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.
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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, respective ly.
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, t he 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 requ irements of various missions. Th is 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.
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• 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 combin ed 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 m 2), 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
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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 wh ich 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.
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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.
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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 high ly 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-ch ip" 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
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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.
Streamlin ed testing is needed for up to 100 or 1000 nanosats per mission. Performing
a complete test program on each unit wou ld 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 miniatu rized 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 yield ing 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 requi res 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 fa ilures.
The large number of nanosats in a constellation is a configuration control challenge for
the data tracking, the schedu les, 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
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associated with a particular nanosat. Constellations that fly in close formation can
benefit by the use of inter-nanosat commun ications to reduce ground station
contention. The data wou ld 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 cou ld 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
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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 fl uid (hydrogen, ammonia,
etc.) then produces thrust by expansion throug h 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 th is 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 liq uid, 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.
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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
12
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plasma, creating downward thrust expansion. Multiple laser pulses and an atmospheric
refresh of breakdown air generate the flight. Th is 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 machin ed 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 lim ited 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.
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Figure 2. AFRL Test Vehicle in Vertical Flight (courtesy of F. Mead, AFRL/PRSP, Edwards
AFB, CA).
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Figure 3. Time-Lapse Photo of a Lightcraft Undergoing an Outdoor Vertical Flight Test
(courtesy of F. Mead, AFRL/PRSP, Edwards AFB, CA).
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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).
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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
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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 con ically
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"
with in which air detonation or propellant ablation (by intense laser heating) occurs; and
3) a parabola-shaped "after body" whose mirrored surface focuses beamed laser energy
into regions of sufficient sma llness for intense air or propellant heating to occur. And as
shown in Figure 8, the vehicle is powered by laser airbreathing propu lsion (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 wh ich 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
of Airflow during
\ • Laser Rocket Flight from
Hypersonic to Orbital
Atmospheric Flight Axi-Symmetric Body Speed
Figure 7. Lightcraft Concept 126 1.
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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 airbreath ing thrusting acceleration during
atmospheric fl ight. 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.
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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)
TakeoffMass=8.0kg ~
FnmlalArea •.096m' ~
I
Takeoff Mass = 8.0 kg
Frontal Area "' .096 m2
(1)
1.00
.16~
i! .0054
(I )
~
Cr,A/W (m21k&)
f
i...= ......s
.10
L .0026 ~ I
·; ...
,II I
~
• I
~41 '-' ~
.. I
.•. l·;...
<lie .. I
·; ..
Q
... i
.01 Range for
Currently l /
/
... -
i1:1 ~ Envisioned ;;iJt' SDIO Llgbtcraft
@ &:ramjets - .,,
~
M
!--
.001
0 10 20 30
Half-Apell Angle ofForebody (degrees)
Figure 9. Lightcraft Vehicle Evolution (in 3 steps) 126 1.
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
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deemed possible with emerging MEMS technologies being developed under the National
Nanotechnology Initiative for both chemica l 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 spread ing - 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," turbu lence, 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 add itional 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.
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~
0.5 • Earth-to-Orbit • End of Luer
Trajectory Alrbreatbiag
\ and Rocket
\ Propulsion
0.4
\
\
\ Beam Power Attenuation
\. j
0.3 \ due to Difraction
Beam Power Attenuation
0.2 0... due to Difraction and
', Atmospheric Effects
,.....,
0.1 ..............
............
"o
0
0 1 :Z 3 4 5 6 7
Laser Waveleogtll/Optimum Laser Wavelength
Figure 10. Attenuation Effects on Captured Laser Beam Power [ 26 1.
0.4 • Earth-to-Orbit
Trajectory
0.3
End of Laser
Airbreathing
Propulsion
o.:z
End of Laser
Airbreathiog
0.1 and Rocket
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.
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Figure 12 shows, for a given laser aperture diameter, adaptive optics, and atmospheric
cond itions, 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-poin ting 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
.:=
f
11.2 micron
Laser Wavelength
La er Beam
Angle from
J:! the Vertical
-:i 1.E+06-l--½---
:i ~a:----=-----F""-..::-------:c:-+.,-------::----:---+------t----t
£
l!
Q,
U
._
1.E+OS-1----- ~ - - - , - - -+-:,--------:----'=l'--..-.:::: - - -- f - -
i...
~
,:! l .E+OA+----~e-------1-----=---1-----------1
0 100 200 300 400 500
Ligbtcraft Slant Range from Laser (km)
Figure 12. Captured Laser Power vs. Increasing Range from 11.2 µm CO2 Laser 126 1.
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 t he 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 rad iated
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. Th us, 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.
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6
•Radiated Laser
1
== s
50• 60°
70°
Power= 10 MW Laser Beam
Angle from
.. 75" the Vertical
-
I
i
~ I I
I
-e.o= 4 IEnd ofLaser1
~ (Airbreathingl
>, 1.62 ·micron
,l:l Propulsion :
Laser Wavelength
] 3
i
I I
-::I
u= 2
Q.
I
I
..
41
I
I • Ligbtcraft
I Capture Dia.
Lightcraftll
Achievement of:
t.. 1 of30 cm Orbital Speed I
=
~
..
GI 11.2 micron
Laser Wavelength
:
\ I
.04
MW
I __j
0
100 200 300 400 soo
Ligbtcraft Slant Range from Laser (km)
Figure 13. Influence of Lightcraft Range and Pointing Ang les on Captured Power C26 1.
The phys ics and technology of FELs will allow beam power to be sca led 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 techno logy that show
promise for achieving megawatt-class beam power include bulk slab solid-state and
high -power fiber lasers; t he former has already achieved over 100 kW of beam power
whi le the latter is getting close to it. Present megawatt-class lasers that are based on
avai lable proven technology include a proposal fo r a 5-beam, 2.5 MW per beam,
electron gun-driven CO2/gas mixture laser which combines five laser beams to ach ieve
10 MW of total beam output power. These systems wi ll 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 Li ghtcraft
ve hicle 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 t he system acq uisition 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
ope rations costs are reduced to one-half those va lues estimated from historical data.
Launch costs are seen to be extremely low (only $74,141 per flight) with laser-
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associated costs comprising approximately 92% of the laser-powered Lightcraft ETO
transportation system LCC.
Table 1. Laser Lightcraft Model Cost Summary [ 26 1.
Laser Lightcraft Model Cost Summary Share Solid-State
Laser Cut Ops
Costs by 50%
Mission Model Length (Years) 10
Launch Rate Per Year 1,000
Payload Per Launch (kg) 2 .0
Mission Flight Time (s) 221.52
Total Program Cost ($M) 741.410
DDT & E / Acquisition Costs ($M) 680 .358 91.77%
Operations Costs ($M) 61.053 8.23%
Average Cost Per Flight ($) 74,141
Average Cost Per kg (based on operations costs) ($) 3,052
DDT & E / Acquisition Costs ($M) 680.358
Laser Lightcraft (LCC) Development Cost ($M) 18.000 2.65%
Laser Lightcraft (LCC) Acquisition ($M) 37.596 5.53%
10 MW Ground-Based Laser Acquisition ($M) 624.762 91.83%
Launch Site Facil ity Costs (Construction) ($M) 5.000 0.73%
Operations Costs, Annual ($M) 6.105
Laser Annual Operations Cost ($M) 3.750 61.42%
Laser Refurbishment, Annual ($M) 0.750 12.28%
Laser Consumables, Annual ($M) 0.00%
Energy Cost, Annual ($M) 0.101 1.65%
Launch Site Facil ity Cost, Annual ($M) 0.250 4.09%
USAF Sys Prgrm Office (SPO) Cost, Annual ($M) 0.250 4.09%
NORAAD Coordination Cost, Annual ($M) 0.500 8 .19%
FAA Coordination Cost, Annual ($M) 0.250 4.09%
Range (Safety, Tracking, Telemetry), Annual ($M) 0.255 4.17%
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Chapter 3: Laser Lightcraft Weapon Mission Selection
Study
The objective of the Froning and Davis [26] 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/sea-based lasers. It entai led 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. See Figure 14 and Figure 15 for
Lightcraft missions that have been identified and explored.
It was concluded that the most promising Air Force mission for a laser-propel led
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 1/S th to 1/10 th the cost required for placing a similar Lightsat
system into LEO using multistage chemical rocket systems.
Ground-to-Space
(Earth-to-Orbit) /
,,,~
/
/
/
Pico or Nano
Satellites
Figure 14. Ground/Sea-to-Space Concept: Appropriate rotation of a high-energy laser beam,
emanating from a ground/sea-based laser, guides and propels an integral Lightcraft pico
/nano-satellite along an ETO ascent path until orbital conditions are reached, or until
beamed laser energy can no longer be transformed into Lightcraft thrust C26 1,
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Enemy/
Air-to-Space Ballistic Missiles
(Ballistic Threats) JI
Air Launch:
H~12 km
/
V~M0.8 / /
Figure 15. Air-to-Space Concept: Appropriate rotation and translation of a high-energy laser
beam, emanating from a moving aircraft, guides and propels a Lightcraft toward a nearly
head-on collision with an incoming ballistic missile; intermittently, the laser illuminates the
target for guidance updates and for terminal semi-active seeker homing and end-game
maneuvering C26 l.
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/sea-based lasers, and no more
than about 2.0 MW for the much lighter and smaller airborne lasers installable on
Boeing 74 7, 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/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
Identified trajectory and mission limitations l261 are:
• Small allowable angle between centerlines of the laser beam and Lightcraft
vehicle axes before significant thrust reduction occurs.
• Small allowable angles-of-attack (angle between Lightcraft centerline and velocity
vector) before significant thrust reduction occurs.
• Limited capability for engaging multiple missile or aircraft threats in allowable
time interval because of relatively long laser beam-riding time against each
threat.
Additional Lightcraft hardware needed for some tactical missions (26 1 include:
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• Termina l seeker guidance for hit-to-kill accuracy against missiles and
maneuvering aircraft.
• Axial-/lateral-propulsion and control for missile and maneuvering aircraft
interception.
• Additional mass along Lightcraft centerline for hardened target penetration.
LIGHTCRAFT MISSION INVESTIGATIONS: SUMMARY AND
CONCLUSIONS
Ground/Sea to Space (ETO) t 26 l
If laser propulsion can provide nearly all Lightcraft /lv needed to reach LEO, then
Lightcraft nanosat systems, which combine both launch vehicle and nanosat
subsystems within a single vehicle, may be achievab le with launch masses in the 2 kg
to 10 kg ra nge; and such Lightcraft nanosat systems appear capable of reaching LEO at
1/S th to 1/10th the cost required using multistage chemical rocket systems.
Air to Space (Air to Orbit) t26 l
If hypersonic magnetohydrodynamic (MHD) airbreathing propulsion research and
development currently underway at the NSF, NASA and the AFRL (Dayton, OH) is
successful, then Lig htcraft 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 MHD airbreathing engines.
Air to Space (Ballistic Missile Interception) t 26 J
Sufficient impact energy for destruction of high-speed ballistic missiles above the
atmosphere is possible with chemical propulsion and uncooled IR detectors (for semi
active homing and axial/lateral acceleration during end-game) integrated into Lightcraft
vehicles for an approximate 100% dry mass increase (from 1.0 kg to 2.0 kg). But
multiple target interception within allowable time is limited by 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 Lig htcraft launches are much
greater. Also, clouds impair Lig htcraft 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/sea-based laser costs are significantly less than multistage chemical rocket
costs. Thus, the selected Lightcraft missions are launch vehicle missions involving
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ground, sea and air launches of Lightcraft to LEO with air launch occurring at either
subsonic or hyperson ic speed.
Lightcraft Ground/Sea to Space Investigation
The current Air Force Lightcraft vehicle concept has been designed for not only placing
nanosats into LEO at low cost (Figure 14 ), but also for performing much of the nanosat
function as well. In this concept the precision optics system that focuses ground/sea
based laser light into t he 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 conventiona l 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 conventiona l rockets was made by Froning and Davis [26]
to get some idea of their comparative costs. Hybrid rocket sizing and costing was
based upon tactical strategic missile sizing and costing information possessed by H. D.
Froning . Th is information related costs (in 1982 dollars) to rocket and payload
characteristics. Lightcraft sizing assumed a propellant mass fraction of 0.5 and 1.0 MW
of laser power per pou nd of payload (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 (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 detail ed future Lightcraft/conventional rocket designs and cost comparisons are,
of course, needed before a strong argument can be made for either design.
Shown in Table 2 is the estimated performance and weights (masses) for 3-stage
hybrid rocket launch vehicles and single-stage laser-powered launch vehicles that are
capable of placing nanosats of 1.0 kg, 5.0 kg, and 10 kg into LEO. And Table 3 shows
estimated costs for hybrid rockets and laser-powered Lightcraft assuming 100 flights
over a 10-year period. These estimated costs indicate that Lightcraft could boost
nanosats in the 2.0 kg to 5.0 kg range into LEO at about 1/l0 t h to 1/5th the cost of
expendable rockets. But Lightcraft cost superiority over conventional rockets is less
overwhelming for satellites that are significantly heavier. Thus Lig htcraft appear
extremely attractive for satellite delivery missions only if Lightcraft dry masses,
including the satell ites being carried, are less than about 5.0 kg.
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Table 2. Performance and Estimated Weights for a Hybrid Rocket and
Lightcraft [ 26 1,
Hybrid Rocket E1pendible Launch Vehicle - Assumed Performance and Estimated Weights
Payload Mass (kg) 1.00 5.00 10.0
3 rd Stage Wt. (kg) 36.8 71.2 108
3rd Stage Isp (s) 295 295 295
Propellant Fraction .737 .766 .782
Impulsive Velocity (km/s) 3.62 3.62 3.62
2".. Stage Wt. (kg) 283 311 341
211c1 Stage lsp (s) 300 300 300
Propellant Fraction .800 .814 .830
Impulsive Velocity (km/s) 3.58 3.58 3.58
1" Stage Wt. (kg) 394 427 462
1" Stage lsp (s) 305 305 305
Propellant Fraction .860 .868 .891
Impulsive Velocity (km/s) 1.92 1.92 1.92
Total Velocity (km/s) 9.12 9.12 9.12
Total Weight (kg) 715 815 921
Lightcraft Expendable Launch Vehicle-Assumed Performance and Estimated Weights
Payload Mass (kg) 1.00 5.00 10.0
Stage Wt (kg) 2.00 10.0 20.0
Effective Isp (s) 1,452 1,452 1,452
Propellant Fraction .500 .500 .500
lmpusive Velocity (km/s) 9.84 9.84 9.84
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Table 3. Estimated Costs for Hybrid Rocket and Lightcraft Launch Vehicles for
ETO Flight l 26 l.
100 Flights
3-Stage Hybrid Rocket 1.0 kg 5.0 kg 10 kg
Launch Vehicle Payload Payload Payload
Payload($) 24,948 124,740 250,128
Boosters ($) 399,503 440,927 477,131
Integration ($) 96,163 101,441 123,941
Total($) 520,615 667,108 851,200
Lightcraft Vehicle.+ 1.0kg 5.0 kg 10 kg
Ground Based Laser Payload Payload Payload
Lightcraft ($) 24,948 124,740 250,128
Lightcraft Fuel {$) 2,495 12,474 25,013
Laser Power ($) 500 2,500 5,000
Laser Refurb. ($) 3,000 3,000 3,000
Total($) 30,943 142,714 283,141
Rocketcraft Cost
16.82 4.67 3.00
Lightcraft Cost
1982 dollars multiplied by 1.62
(2.7% annual inflation) for 2000 dollars
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Lightcraft Air to Space Investigation
One Ballistic Missile Defense (BMD) mission already being investigated by the Air Force
for high-power airborne laser systems is the focusing of their intense beam energy on
enemy ballistic missiles over dwell times sufficient to heat missile ma
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