Pulsed Laser Thermal Propulsion For Interstellar
Precursor Missions
Jordin T. Kare
Kare Technical Consulting, 222 Canyon Lakes Place, San Ramon, CA 94583
(925) 735-8012; jtkare@ibm.net
Abstract. A laser-thermal propulsion system is proposed for launching large numbers of small interstellar precursor
probes at velocities up to ~300 km/s (0.001 c). This system uses a stationary pulsed laser, based on Earth or in nearEarth space, to beam energy to probe vehicles during their initial acceleration. Each vehicle collects laser energy using
a deployable reflector, and focuses the laser energy into a thruster. The focused laser pulses ablate and heat an inert
propellant, which expands to produce thrust at a selectable specific impulse up to of order 20,000 seconds (exhaust
velocity up to 200 km/s). This technology permits the vehicles to be simple and light, while allowing much higher
acceleration than alternative propulsion systems. The laser system is ideal for launching large numbers of flyby probes,
for example to examine many objects in the Oort cloud. A laser system with 30-meter-class transmitting optics and a
100-MW laser is capable of launching 100 kg payloads to 50 km/s, with payload mass fraction (probe payload / probe
initial mass in Earth orbit) of 10-20%. The same system can launch much larger payloads to lower velocities for Solar
System exploration. Scaling relationships are derived and scaling options discussed, along with possible near-term
development and proof-of-concept tests.
INTRODUCTION
Interstellar precursor missions such as Kuiper object flybys, Solar Focus missions, and the TAU (Thousand
Astronomical Unit) mission involve launching spacecraft to ranges of 50 – 1000 AU. To complete these missions
in reasonable time (5 – 20 years) requires velocities of 10 – 50 AU per year, or 50 – 300 km/second mean mission
velocity.
Current trends in miniaturization are shifting mission designs toward small spacecraft (100 – 1000 kg) and
microspacecraft (<100 kg). For many missions, such as exploration of the Kuiper belt and Oort cloud, large
numbers of such small vehicles may be inherently more productive than single large spacecraft, since no single
“target” exists. A high ∆V propulsion system matched to small vehicles would thus be highly desirable, especially
if it permitted launching many missions at low per-mission cost.
PROPULSION CONCEPT
Laser propulsion is a beamed-energy propulsion concept in which a stationary ground- or space-based laser
transmits energy to the spacecraft. Because the expensive components of the system – the laser and transmitting
optics -- remain stationary, they can be reused over many missions. Laser-electric propulsion, in which the laser
energy is converted via photovoltaic (PV) cells to electricity to drive an electric thruster, has been evaluated for
both near-Earth and interstellar missions (Forward 1985, Jackson 1978, Landis 1989), but has limited thrust-toweight and high per-mission cost due to the large PV cell area needed.
In laser thermal propulsion, the laser energy is used directly to heat a propellant, which is exhausted to produce
thrust. Laser thermal propulsion has been considered mainly for Earth launch and near-Earth maneuvering, using
either continuous (CW) or pulsed lasers and various types of thruster (Glumb 1984, Kare 1990, Kare 1995, Phipps
1994). The exhaust velocities required for interstellar precursor missions are too high for practical operation of a
CW thruster with a material nozzle. CW thrusters with specific impulse (I sp) >> 1000 s may be possible using
magnetic nozzles, but this author is not aware of any studies of this option.)
Pulsed ablation thrusters use the laser energy to ablate a thin layer of a (usually solid) propellant. If the laser flux
and fluence are sufficiently high, the ablated material will form an absorbing plasma which can be further heated
by the beam to extremely high temperatures. The actual plasma temperature can be controlled by varying the laser
flux and fluence, and to some extent by varying the laser pulse shape.
A pulsed thruster may or may not use a divergent nozzle to constrain the plasma expansion. For sufficiently short
pulses, the plasma is created in a layer that is thin compared to the width of the thruster, and expansion is
effectively one-dimensional over most of the thruster even without a nozzle. Such a nozzleless thruster is
particularly simple, and has the advantage that the laser illumination can come from any angle, while the thrust is
always normal to the propellant surface.
The primary components of a laser thermal propulsion system for in-space use are shown in Figure 1. The laser
produces an average power P, and may be continuous (CW) or pulsed. The laser beam is expanded through
various optics and finally focused by a transmitting aperture of diameter Dt, which may be a large diffractive
element or concave mirror. We assume for this paper that the beam leaving the transmitting aperture is diffractionlimited. The beam is received at the vehicle through by a collector of diameter Dc, which may also be a diffractive
lens or mirror, and can generally be of much poorer accuracy. The collector focuses the beam onto (or into) the
thruster.
Laser power
supply
Transmitter optics
Collector
Space-based laser
Dt
Dc
Space relay optics
Dt
Ground-based laser
Adaptive optics/
beam director
Payload
v ex
v
Laser-driven thruster
FIGURE 1. Laser Propulsion System Concepts for In-Space Propulsion.
Pulsed Ablation Thruster Characteristics
A laser pulse with energy E at the thruster ablates a mass m of propellant. In vacuum, the resulting plasma
expands, reacting against the remaining propellant and nozzle, if any, and eventually reaching a mean directed
exhaust velocity v ex relative to the vehicle. By conservation of momentum, this results in an impulse of m * v ex ,
and a specific impulse of v ex /g, where g is 9.8 m/s2. vex depends on the material ablated, the laser flux and
fluence, and to some extent the laser wavelength and pulse shape.
The thruster efficiency thr is defined as:
thr = m v ex 2 / 2E,
(1)
and represents the fraction of laser energy that is coupled into useful exhaust kinetic energy (as opposed to
radiation, lateral motion, or internal energy of the exhausted gas). Thruster efficiency is also dependent on the
ablated material, pulse properties, etc. Simple ablation of common materials (aluminum, plastics, etc.) can produce
exhaust velocities of 103 to 105 m/s with efficiencies of 10 – 50%, although with considerable scatter in the data
(Phipps 1994). With optimized material and pulse properties, we assume th r = 0.2 will be consistently achievable
up to v ex = 200 km/s While this efficiency may seem low, it is comparable to the efficiency of a laser-electric
propulsion system:
(PV cell) x (power conversion) x (electric thruster) ≈ 0.4 x 0.95 x 0.5 = 0.19.
(2)
For convenience, we define an exhaust power Pex = th r opt P, where opt is the optical transmission efficiency
of the laser beam, including diffraction, scattering and reflection losses, etc.
SYSTEM SCALING
The laser propulsion thruster exhaust velocity can be chosen by the system designer; for a given laser power,
higher v ex yields lower thrust (as 1/ v ex , assuming fixed th r ) and lower propellant consumption (as 1/ v ex 2).
Assuming both P and v ex are held constant over the acceleration period, the mass ratio and relative time to
accelerate a given final mass mf to a final velocity vf varies with v ex as shown in Figure 2. Acceleration time (and
therefore total laser energy used) are minimized at vf / v ex ~1.63, but varies slowly for 1 < vf / v ex < 2.
Starting from rest, the range at “burnout” is given by:
R
 e    1,
m f v f 3 3

2 Pex
(3)
where  = vf / v ex . The variation in range with vf / v ex is also shown in Figure 2; the minimum range occurs at
slightly lower exhaust velocity and higher mass ratios than the minimum time. Slightly shorter times can be
achieved by varying the exhaust velocity over time to approximate v ex (t) = v(t) (which gives maximum energy
efficiency) at the expense of substantially higher mass ratios. Actual systems may choose to use higher-thanoptimum-range exhaust velocities to minimize propellant mass launched from Earth, although the mass ratios
involved are small by chemical rocket standards. We assume in scaling estimates below that vf / v ex = 1, so that the
FIGURE 2: Variation of acceleration (burn) time, range, and mass ratio with vf / v ex .
final mass is ~35% of the initial mass in Earth orbit.
The range at burnout is limited by diffraction of the transmitted beam. This limit is given by:
Burn time, range (arb. units)
Rm ax 
Dt Dc
.
fop t
(4)
of range. For simplicity, we take fopt = 2 and fix opt at 0.5 in the system calculations below.
Mass ratio
Traditionally, fopt is taken to be 2.44, which corresponds to the collector just spanning the first null of the far-field
Airy pattern for an initially uniform beam and collecting 84% of the beam energy. In fact, a useful fraction of the
transmitted beam power can be collected at substantially longer range, and opt should be calculated as a function
(The effective value of Rmax can be increased – in the limit, doubled – by starting acceleration “uprange” of the
transmitter aperture, rather than starting from the vicinity of the transmitter. This requires that the transmitter be
able to track the vehicle over a large angle at a significant angular rate, and that the vehicle be able to accept a
beam from in front as well as behind. This factor is not included in the system performance calculations.)
Setting R = Rmax and rearranging, we get:
 D D  2P th r op t 
m f   t c  


,
3
 fop t    v f f (  ) 

(5)

where f (  )   3 e     1 ≈ 0.52 (minimum) to 0.72 ( =1). All of the terms on the right are independent
variables except the collector diameter Dc . As the final mass increases, clearly the mass available for the collector
also increases. Assuming a fixed fraction f c of mf is allocated for a collector (and its support structure) with a
mean areal density :
fc m f 

4
 Dc 2 or Dc 
4 fc m f

.
(6)
Defining a “payload” mass (vehicle dry mass excluding collector) m pa y  1  fc  m f , we have:
2
2
4 f   D  2 P th r op t 
m pa y  1  fc   c   t  
 v f 6 ,
f ( )
    fop t  

(7)
which is obviously maximized at f c = 0.5, i.e., roughly half the mass remaining at burnout should be collector.
A variation of v f  6 is quite impressive, suggesting that, for example, a system capable of launching 1 kg to 300
km/s would at least in principle be able to launch 1000 tons to 30 km/s. In practice, collector mass is likely to vary
somewhat faster than Dc 2, and eventually reach a practical limit for a given technology, but mf will still vary
rapidly with v f .
Dt is limited by cost and available space optics technology. In general, telescope apertures vary in cost
approximately as D2.5. Assuming P is proportional to the laser cost, it can be shown that the optimum allocation of
cost is 2/7 to the transmitter aperture and 5/7 to the laser, and in this simple model mf will vary approximately as
the system capital cost to the 2.8 power, or, for fixed payload mass, the system cost will vary roughly as v f 2.
PROJECTED SYSTEM CHARACTERISTICS
Existing high average power lasers operate primarily in the range between 1 and 10 µm. Future high power lasers
may be expected to operate between 0.3 and 1 µm. Current costs for ground-based high-average-power lasers are
well over $1000/watt (Phipps 1996). However, both diode-pumped solid state lasers and free-electron lasers have
the potential to produce power levels of 10 – 100 MW at ~$100/watt or less. Reaching $10/watt, even using
ground-based laser hardware, will require significant progress or breakthroughs such as low-cost phase-locking of
large arrays of diode lasers, but seems plausible in 20 to 30 years.
Space optical systems for the near-IR planned for the next 10 years, such as the 8-meter Next Generation Space
Telescope, have masses of ~15-20 kg/m2 (Stockman 1997); advanced reflective optics technologies are under
development with predicted masses ranging from 10 to 1 kg/m2. Diffractive optics, which are well suited for use
with monochromatic light, are proposed with aperture masses well below 1 kg/m2 at 25-50 m aperture diameter
(Hyde 1998). We can therefore reasonably expect a 10,000-kg-class, $1-B class optical aperture to evolve from
~10 to ~100+ m diameter over the next few decades.
Collector optics may be extrapolated to a small multiple of the mass of a reflective film, assuming either a very
lightweight (e.g., gas-inflated) concave reflector or an embossed holographic concentrator. In the former case, the
short operating life may allow lighter structure than would be required for, e.g., a long-lived solar concentrator.
With current technology, inflatable solar concentrators have areal densities of <0.2 kg/m2 (5 m diameter, 3.3 kg
(SRS Technologies 1999)). An aggressive collector technology, using, e.g., embossed aluminum film 1 µm thick
as a holographic reflector, could reduce this by an order of magnitude, with corresponding tenfold increase in
payload. This is still several times the areal density proposed for aggressive solar or laser sail designs (Matloff
1984).
Figures 3a and 3b shows the laser power and transmitter aperture diameter for various mission velocities for a
nominal mission with mpay = 100 kg. Figures 3c and 3d give the system range and the time to accelerate one
payload. Note that these are still “first approximation” values, assuming free-space acceleration, and do not
account for (or take advantage of) orbital dynamics. Figure 3e shows a calculated “system cost” based on simple
scaling of the laser and telescope costs only. This cost is obviously only a rough indicator of what a real system
would cost, but does indicate how the cost scales with mission velocity and technological progress. The
assumptions used to derive Figure 3 are given in Table 1.
TABLE 1. Nominal Parameters for Laser Propulsion System Scaling.
Parameter
Near term (2010)
Mid term (2020)
Far term (2030)
Laser cost ($/watt)
100
30
10
Telescope cost
10
30
100
baseline (Dia. of $1B
telescope, m)
Collector areal
0.2
0.1
0.02
density  (kg/m2)
Wavelength (µm)
1
0.5
0.33
Common parameters: th r = 0.2, opt = 0.5, fopt = 2,  =1, payload mass 100 kg.
Several features stand out in Figure 3. First, laser power levels and transmitting apertures for interstellar precursor
missions, while large by current standards, are well within reach of foreseeable engineering. Lasers with power
levels of 100 MW to 1 GW have been proposed (and analyzed in some detail) for strategic defense. 100-meterclass optics are well beyond the curent state of the art, but are plausible with thin-film mirrors or diffractive optics,
especially for a narrow-band, narrow-field-of-view system. Second, a laser propulsion system for interstellar
precursor missions benefits strongly and synergistically from progress in several technology areas, notably lasers,
space telescopes, and thin-film collectors. System performance improves essentially as the product of the
performance of these individual technologies, rather than being driven by a single “long pole”. Thus, despite the
extremely steep relationship between mission velocity and system parameters, the mission velocity achievable at a
given level of system cost will almost certainly increase steadily over the next few decades, and in the long run,
laser propulsion is likely to be superior to other propulsion approaches which are dependent on progress in a single
technology.
FIGURE 3. Variation in laser propulsion system characteristics vs. mission velocity, for a 100 kg payload: (a) Laser power, (b)
Transmitter aperture, (c) Range at burnout, (d) Acceleration time, and (e) Estimated basic system (laser + transmitter optics) cost.
Parameters for near-, mid-, and far-term system are given in Table 1.
OPTIONS FOR DEVELOPMENT AND DEMONSTRATION
Near term development of high-Isp pulsed ablative propulsion requires mainly incremental extension of existing
laser-ablation research, including laser lethality studies, to repetitive pulses, and to materials chosen for optimum
propellant properties. Repetitive pulses are needed to measure the efficiency and exhaust velocity in steady state.
Models of laser ablation processes and gas/plasma flow can be used to predict efficiency in operation and
investigate, e.g., required beam uniformity and optimum laser pulse profile.
Demonstration of propulsion in space will require either a laser with high pulse energy and reasonable (kilowattclass) average power in space, or such a laser on the ground with suitable adaptive and relay optics to propagate a
high quality beam to at least low Earth orbit. A plausible “technology demonstration” system would have a 100
kW-class 1 µm laser, presumably ground-based, and 3- to 4-meter transmitting aperture. Such a system has
sufficient optical range (~30,000 km, to an 18 meter, 100-kg collector) to provide high-Isp propulsion anywhere in
sub-geosynchronous space., At a thruster Isp of 1000 s, it would take 14.6 hours of thrusting to accelerate a 100 kg
payload by 1 km/s. A demonstration system could thus provide useful orbital maneuvering propulsion as well as
serving as a testbed for a larger system for interstellar precursor missions.
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