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Photon Tether Formation Flying (PTFF) Bae Institute
A Contamination-Free Ultrahigh Precision Formation Flight
Method Based on Intracavity Photon Thrusters and Tethers:
Photon Tether Formation Flight (PTFF)
Phase I Program 07605-003-041
Phase I Final Report
Prepared for:
Robert A. Cassanova, Ph.D., Director
NASA Institute for Advanced Concepts
April 30, 2006
PI: Young K. Bae, Ph.D.
BAE INSTITUTE
Tether
“I believe in intuitions and inspirations.
I sometimes feel that I am right.
I do not know that I am.”
by Albert Einstein
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CONTRIBUTORS
Young K. Bae, Ph.D. Bae Institute
C. William Larson, Ph.D, Air Force Research Laboratory
T. Precilla, Ph.D. Northrop Grumman
Joseph Carroll Tether Applications Inc.
Claude Phipps, Ph.D. Photonics Associates
Cover: An artist’s concept of a km-diameter ultralarge membrane telescope formed by
spacecraft in nano-precision formation flying (not to scale). A small space telescope
shown left side is an artist’s concept of James Webb Space Telescope for size
comparison. The second panel shows the general force structure of formation flying
and examples of nano-precision structures that can be formed with the proposed
method.
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TABLE OF CONTENTS
COVER ------------------------------------------------------------------------------------------- 1
COLLABORATORS ------------------------------------------------------------------------- 2
I. INTRODUCTION --------------------------------------------------------------------------------- 3
II. PHOTON TETHER FORMATION FLIGHT (PTFF) -------------------------------------- 9
A. General System Concept --------------------------------------------------------------------- 9
B. Photon Thruster System --------------------------------------------------------------------- 12
C. Interferometric Ranging System Integrated with Photon Thruster System -- 22
D. Tether System ------------------------------------------------------------------------------------- 25
III. PTFF TECHNOLOGICAL READINESS ASSESSMENT ------------------------------- 36
IV. MISSION SPECIFIC APPLICATIONS ------------------------------------------------------ 37
A. Low Earth Orbit (LEO) Applications ------------------------------------------------------- 38
B. Geosynchronous Earth Orbit (GEO) Applications ------------------------------------ 40
C. Lagrangian and Other Orbit Applications ----------------------------------------------- 41
V. EXEMPLARY MISSION STUDIES ----------------------------------------------------------- 42
A. Ultralarge Adaptive Membrane Telescopes -------------------------------------------- 42
B. New-World-Imager and Freeway Mission ----------------------------------------------- 44
C. 1-D Formation Flying Structure for Fourier Transform X-Ray (FTXR)
Interferometer ------------------------------------------------------------------------------------ 45
VI. DESIGN OF PHOTON THRUSTER AND NANO-PRECISION THRUSTER TEST
STAND FOR PHASE II ------------------------------------------------------------------------ 49
VII. ROADMAP ---------------------------------------------------------------------------------------- 53
A. Predictions and Limitations on PTFF ----------------------------------------------------- 53
B. Required Technologies ---------------------------------------------------------------------- 55
C. Phase II and Beyond -------------------------------------------------------------------------- 57
VIII. REFERENCES ---------------------------------------------------------------------------------- 59
APPENDIX -- Publication in STAIF 2006 Proceedings ------------------------------- 61
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OVERVIEW
We proposed a revolutionary nano-meter accuracy formation flight method based on photon
thrusters and tethers, Photon Tether Formation Flight (PTFF), with the maximum baseline
distance over 10 km, and have investigated its feasibility in depth during this project. In
addition, PTFF is predicted to be able to provide an angular scanning accuracy of 0.1 micro-
arcsec, and the retargeting slewing accuracy better than 1 micro-arcsec for a 1 km baseline
formation.
The conclusion of this study is that the implementation of the proposed method in the near
future is well within reach of the present technologies. Therefore, we are confident that the
proposed PTFF needs thorough continued study that will establish a reliable technical path to
the launch of an exciting new class of NASA space mission. In this report, we review how we
arrived at the conclusion and the scientific and technological potential of the method, and show
how it can be realized and implemented in numerous NASA space missions. We present some
instrument design concepts for PTFF and analysis of their technological readiness in terms of
the NASA TRL scale. From these we identify which are the key technologies that need attention
before the proposed method is implemented in a mission with confidence.
The core technology of PTFF is in the strategic combination of photon thrusters and tethers,
which provides propellant-free, thus contamination-free, ultrahigh precision control. The
intracavity arrangement was exploited to trap photons between two spacecraft multiplying the
thrust by several orders of magnitudes. As a result, the thrust power requirement for formation
of 100 kg spacecrafts can be only several watts per pair, within the power budget. The
intracavity arrangement further allows ultrahigh precision thrust vector control that cannot be
achieved with the conventional thrusters. Propellant mass saving and longer mission lifetime
(tens of years) are other advantages of the proposed concept.
Furthermore, a portion of ultra-stable photon thruster lasers can be used for nano-meter
accuracy interferometric ranging simplifying the system design and reducing the system weight.
This approach also eliminates the necessity of additional optical delay line system that was
necessary in the existing formation flying methods, thus it results in further weight reduction and
system simplification. In addition, PTFF is capable of providing near adiabatic CW control that
minimizes generation of tether vibration. Tether vibrations potentially caused by meteoroid
impact and major formation structure reorientation, have been addressed and shown to be
readily quenched with electromechanical dampers and photon thruster power modulation.
In addition to redefining and simplifying the existing NASA mission concepts, such as SPECS
and MAXIM, PTFF enables other emerging revolutionary mission concepts, such as New World
Imager Freeway Mission proposed by Prof. Cash, which searches for advanced civilization and
in exo-planets Fourier Transform X-Ray Spectrometer proposed by Dr. Schnopper. As the
present concept is more publicized, many other exciting concepts are expected to be inspired
by PTFF. One of such possible NASA missions is the construction of ultralarge space
telescope with diameters up to several km for observing and monitoring space and earth-bound
activities.
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I. INTRODUCTION
In recent years, microsatellites and nanosatellites enables the insertion of sophisticated
sensors and processing technologies into orbits of interest at low costs (Leitner, 2004)
for the next generation NASA missions. Building a cluster of small satellites has been
recognized to be more affordable, robust and versatile than building a large monolithic
satellite. Specifically, the grouped satellite cluster is crucial for enabling orders-of-
magnitude improvements in resolution and coverage achievable from advanced remote
sensing platforms. Size limitations on launch vehicle fairings leave formation flying as
the only option to assimilate coherent large apertures or large sample collection areas in
space (Leitner, 2004).
Numerous NASA mission applications have been envisioned; for example, the ultrahigh
precision satellite clusters can eable interferometry and distributed large aperture
sensors, especially at optical (TPF, and SPECS) and x-ray wavelengths (MAXIM)
(Leisawitz, 2004, Cash, 2002). For non-NASA applications, the proposed system can
enable advanced geophysical monitoring where GPS and standard laser range finders
are currently inadequate to measure and monitor small changes in the movement of
earthquake plates, and gravity wave detection. Other commercial and military
applications include distributed large aperture optical and infrared sensors for ultrahigh
resolution monitoring and imaging at low-cost.
Such a technology critically depends on the formation flying method that enables
precision spacecraft formation keeping from coarse requirements (relative position
control of any two spacecraft to less than 1 cm, and relative bearing of 1 arcmin over
target range of separations from a few meters to tens of kilometers) to fine requirements
(nanometer relative position control). For example, one of the most challenging
applications for formation flying thus so far is that of the proposed x-ray interferometry
for space imaging applications, MAXIM (Cash, 2000). The concept has evolved to
include a pathfinder mission, consisting of a single x-ray interferometer and a trailing
imaging satellite, and the full MAXIM, consisting of a fleet of 33 x-ray mirror satellites, a
trailing collector satellite, and an imaging or detector spacecraft.
The formation flying baseline accuracy for optical interferometry applications can be
relaxed with the use of optical delay lines (ODL) as has been proposed in the existing
formation flying concepts. However, this concept requires two independent systems,
the spacecraft position control system and optical delay line control system, for
controlling two different types of random motions of which magnitudes are several
orders of magnitude different. Therefore, the control system would be exceedingly
technologically challenging. Another simpler approach to this problem is the use of the
spacecraft platform as optical platforms. Although this approach highly simplifies the
system integration and reduces the weight, it requires the baseline control accuracy that
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are much more stringent than that can be obtained with the conventional control method
with traditional thrusters. In this case, summary of the requirement of the baseline
accuracy tolerance of several exemplary missions compared with the capability of the
present formation flying method is shown in Fig. 1.
10-9 10-8 10-7 10-6 10-5 10-4 10-3 10-2 10-1 Meter
MAXI
M
SPECS TP
F
Present FF Metho
d
FIGURE 1. Required Base Line Accuracy of Several Exemplary NASA Missions
and the Capability of the Present FF Method. The accuracy is estimated for the
system without the optical delay line system.
In MAXIM, the relative distance between the hub satellite and collector satellites should
be precisely maintained with the tolerance of a few nm (10-9 m) at the distance of 200 m,
and the precision requirement in maintaining the distance, thus, is 10 parts per trillion,
one of the most stringent accuracy requirement seen in any scientific fields. In addition,
potential contamination of neighboring spacecraft by propellant exhaust plumes and the
possibility of pulsed electromagnetic interference with low power inter-satellite
communications remain a real concern for grouped satellite clusters. These
requirements essentially rule out the usage of the most of the conventional propellant
based propulsion systems, such as gas hydrazine thrusters, pulsed plasma thrusters,
hall thrusters, electrostatic ion engines, and field emission electron propulsion systems.
To alleviate the second critical concern, the propellant plume contamination, several
propellant-free formation flying methods have been proposed. The propulsive
conducting tethers and spin-stabilized tether systems have been proposed in place of
on-board propulsion systems to form and maintain satellite formations (Johnson, 1998,
Quinn, 2000).
While such concepts offer intriguing possibilities for small arrays
consisting of only a few spacecraft, implementing a system for dozens of satellites
quickly becomes extremely problematic. Several other new concepts have been
proposed. They are: 1) the microwave scattering concept (LaPointe, 2001), 2) Coulomb
force concept (King, 2002), 3) magnetic dipole interaction concept (Miller, 2003). In the
microwave scattering formation flight method (LaPointe, 2001), radiation forces on the
order of 10-9
N/W may be generated using electromagnetic gradient forces or scattering
forces; microwave beam powers of 10-kW can thus produce restoring forces of
approximately 10-µN, which are sufficient to correct a number of orbital perturbations. It
requires very high power consumption, and focusing of the microwave requires larger
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antenna arrays, and the scattered microwaves may electronically interfere with other
neighboring satellites. The Coulomb control system (King, 2002) is limited to close
formation flying in plasma environments characterized by Debye lengths greater than
inter-vehicle separation. Even for such formations, however, the Coulomb control
forces become negligible for separations greater than 50 m. It is apparent that more
traditional thrusters would be necessary for formation keeping over larger distances.
Generating usable Coulomb control forces requires charging spacecraft to high
voltages, thus great care must be taken in vehicle design to prevent differential charging
and instrument damage due to electrostatic discharge. In the magnetic dipole concept
(Miller, 2003), two technical challenges should be overcome: 1) it may not work at the
distances greater than tens of meters, thus cannot be used for SPECS and MAXIM, 2)
the system can be extremely bulky and heavy (tons). Therefore, searches continue for
the concept that does not require propellant nor extreme high voltages, and is power
efficient and light.
Even if an efficient propellant-free thrust is developed, for ultrahigh precision formation
flying, the issue of thrust pointing, the method of controlling thrust to the desired
accuracy, the precision ranging metrology and the overall system architecture should be
addressed. Thus far, a solution for maintaining a precise spacecraft configuration that
can be used as an optical platform in space has proven illusive.
We have proposed and investigated under this NIAC Phase I contract the feasibility of a
potentially revolutionary formation flying method, Photon Tether Formation Flight
(PTFF), which enables ultrahigh precision spacecraft/satellite formation flying with
intersatellite distance accuracy of nm (10-9 m) at maximum estimated distances in the
order of tens of km in principle. The method is based on an innovative ultrahigh
precision intracavity photon thruster able to provide continuously adjustable thrust from
pN (10-12 N) to mN between spacecraft/satellites, and tethers. The thrust of the photons
are amplified by as much as 20,000 times by bouncing them between two mirrors
located separately between pairing satellites, and a 10 W photon thruster, which is
suitable for microsatellite formation flying, is capable of providing thrusts up to 1.34 mN
with currently available components. This thruster efficiency well rivals that of the most
efficient electric propulsion system. A crystalline-like structure of satellites is proposed
to be formed by the pushing-out force of the intracavity lasers and the pulling-in force of
the tether tension between satellites.
The salient features of the present concept are:
• The proposed photon thruster system is capable of generating propellant-free
continuously for tens of years,
• In combination of a tether system, the proposed photon thruster is capable of
controlling and maintaining continuously the intersatellite distance with an
accuracy better than nanometer,
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• The dual usage of the photon thruster as a laser source for the ultrahigh
precision interferometric ranging system simplifies the system architecture and
control, and minimizes the system weight and power consumption,
• The thrust vector can be defined with ultrahigh precision due to the nature of the
laser cavity.
• The method can be readily scaled down to nano- and pico-satellite formation
flying.
Ever since the concept was proposed, though there seems to be no violation of physical
principles in the concept, numerous daunting engineering challenges have emerged. It
seems that more engineering problems will be identified as the program progresses.
The challenging engineering problems identified so far can be classified into three
categories:
1. The photon thruster system issues
• The practical limit on operation distance of the proposed concept
• The practical limit on the reflectance of the mirrors
• Thermal management of the laser system, in particular, that of the laser crystal in
the cavity
• The method to provide maximum dynamic range with maximum distance
precision
• The operational lifetime of the laser system and how to extend it
• The method of the pointing system to align the laser beam
1. The tether system issues
• The hardware to control/adjust the length
• The method to damp vibrations or resonances in the tethers
• The method to adjust the thermal expansion/contraction
• The lifetime of the tethers in the space environment
• The method of coping the breakage of tethers due to micrometeoroid impact
2. Coordination of photon thrusters with tethers and overall system control issues
• The method to compensate various perturbations, such as the solar pressure,
drag force, thermal expansion/contraction, in the space environment
• The method to obtain ultrahigh precision relative position information of satellites
• The engineering architecture of the multiple satellite system
• The control/command method
During this Phase I NIAC program, we have investigated these issues, and concluded
that the proposed concept is indeed feasible. Many of these technical and engineering
issues have been tackled during Phase I study, and we plan to continue this effort in
Phase II. The following sections summarize the results of the study.
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II. PHOTON TETHER FORMATION FLIGHT (PTFF)
A. General System Concept
The Photon Tether Formation Flight (PTFF) enables ultrahigh precision satellite
formation flying with intersatellite distance accuracy of nm (10-9 m) at maximum
estimated distances in the order of tens of km. Thus, the present method can be used
for most of next generation NASA formation flying missions envisioned so far, including
TPF, SPECS, and MAXIM without the complication of the optical delay lines. The
method is based on innovative ultrahigh precision laser intracavity thrusters able to
provide continuously adjustable precision CW thrust between microsatellites and tethers.
The present concept can be used for both spinning and non-spinning systems. In
slowly spinning systems, as in NASA SPECS, centrifugal force can provide a precise
repulsive force, allowing a low-mass tether to provide precise control of distance. There
is a problem how quickly that force can be adjusted without risk of inducing undesired
resonances, and to solve the problem, the agile control loop can use adjustable laser
power rather than mechanical tether length control as the primary control mechanism. In
non-spinning systems centrifugal force is not available, and a laser will provide the
major repulsive force. In both spinning and non-spinning cases, the fast feedback
possible can be used not just to control position, but to reduce the required agility of the
tether control, and hence the problems induced by undesired tether resonances.
The schematic diagram of the proposed concept is shown in Fig. 2. Specifically, the
proposed formation flying method is based on pulling-in force provided by tether tension
and the pushing-apart CW thrust of the intracavity photon thruster. Although the thrust
produced by single bounces of photons is typically negligibly small, the intracavity
geometry allows photons to bounce between two mirrors as many times as tens of
thousands, resulting in several orders of magnitude amplification of the thrust with a
given laser power. With this proposed method, we estimate the distance between the
satellite pairs in the constellation structure can be adjusted and maintained rapidly to
the accuracy of nanometer. The photon thrust and tension of tethers form the backbone
linear force structure of the crystalline-like structured formation flying, and can rapidly
damp the perturbation from the space environmental sources, such as solar pressure,
drag-force, and temperature fluctuation, applied from any direction.
Several exemplary structures and the general force structure are illustrated in Fig. 2.
For example, the tetrahedral structural can be used for NASA SPECS applications. As
illustrated in Fig. 2, the present concept “inflate” the ultralarge space structure with
bouncing photons, and the structure boundary is limited and stabilized by tethers. For
NASA MAXIM applications, an elongated polygon bipyramidal structure similar to the 10
satellites constellation can be used, except instead of 8 satellites 32 collector
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spacecrafts will be used in the plane. The apex will be occupied by the hub and
converger crafts. The approximate distance between the collector and hub crafts is
about 100 m, and that between the collector and the converger crafts is about 10 km.
This is slightly modified structure from the original one (Cash, 2000). These numbers
will be used to estimate the necessary operation parameters in the later sections. For
one to two dimensional structures, such as in ST-3 and space Fourier Transform X-Ray
Spectrometer, recently proposed by Dr. Schnoepper, multiple photon thrusters and
tethers are necessary for a pair of satellites to stabilize the angular disturbance.
Figure 2. Schematic Diagrams of the Exemplary Satellite Mission Configurations
with the Proposed Ultrahigh Precision Formation Flying Method. The 8-Satellite
Constellation may require additional triangulation for stability.
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More specifically, for SPECS applications, the proposed concept here has two important
advantages. First, the usage of tethers will obviate the need for a massive amount of
thruster propellant and high-quality imaging interferometry with a 1 km maximum
baseline. "High quality imaging" implies the need for dense coverage of the "u-v" plane
(i.e., moving the light collecting telescopes to fill the area subtended by the synthetic
aperture). To produce images at a reasonable rate, the light collectors will have to be
moved around a lot. Although even the most efficient thrusters available can't perform
such a task, the tension in a tether can do nearly all the work. Second, to operate with
the required sensitivity at far-IR wavelengths, SPECS will have cryogenic optics
maintained at 4 K. Therefore, contamination of the optical surfaces is a big concern.
The proposed system can be used as an alternative version for the originally proposed
for SPECS to overcome these concerns.
The more detailed schematic diagram of the system architecture of the proposed
system is shown in Fig. 3.
Satellite I Satellite II
Precisio n Laser
Po wer Me te r
HR Mi rror HR Mirror
Las er G ain M edia
Ultrahigh Precision C W Photon Thrus t
Tet her
Tension
Piezo-Translator
Stepper Motor
Intracavity Laser Beam
Te th e r
Reel
Clamp
Lens
Diode
Pump
Laser
Pump
La ser
Beam
HR Mirror
HR Mirror
Part al Mirrori
Photodetector
Part a l Mirrori
Part al Mirrori
Tether System
Photon Thruster
System
Interferometric
Ranging System
Figure 3. The schematic diagram of the combined system of photon thruster,
interferometric ranging and tether systems. The electromechanical damper is
configured with the tether clamp, and will be able to efficiently damp the
transverse tether vibration efficiently.
The system has three major sub-systems: 1) the photon thruster system, 2) the
interferometric ranging system, and 3) the tether system. The proposed ultrahigh
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precision photon thruster will provide the thrust that will push the satellites apart
resulting in extension as a function of the laser thrust. The opposite force is balanced
by the tether tension and the length of the tether is proposed to be adjusted by linear
translators composed of piezoelectric translators and stepper motors to the accuracy
better than 1 nm. The ultrafine balance of the pulling-in and pushing-apart is
maintained by the above two mechanisms and controlled by a computer in real time,
forming crystal structure-like satellite cluster. The vibration perturbation or resonance
induced by the actuation and motion of the systems and satellites in tether can be
rapidly (almost in real time) damped by the photon thruster and electromechanical
damper that is not shown in Fig. 3, but will be described in detail in the later sections.
For example, diode pumped solid state laser systems, such as a diode pumped YAG
laser intracavity laser system, can be used for the proposed formation flying method. In
this case, we estimate that the extracavity laser power in the order of 10 W capable of
providing photon thrust up to mN is suitable for the weight of each satellite in the order
of 100 kg. The power consumption and weight of such a laser system are estimated to
be about 30 W and several kg, respectively. The intracavity laser beam is formed
between two high reflectance (HR) mirrors located in two separate satellites. The
matching tether diameter in this case, for example, is in the order of 4 mm with Kevlar
fibers. The power of the laser and the inverse of the cross sectional area of the tether is
linearly proportional to the weight of the satellite. For example, the formation of 10 kg
and 1 kg satellite constellations, the required laser powers are in the order of 1 W and
0.1 W respectively. The weight of the laser system decreases rapidly as the laser
power decreases, thus the technology can be easily adapted to much smaller and
lighter satellite platforms. Eventually, the limiting factor in weigh scaling down will be
the weight of the tethers and associated structures.
Another important issue that the effect results from that the spacecraft has size,
moments of inertia, and attitude perturbation torques. Adjusting moving the tether
attachment points may be the best way to get rid of steady-state "bias torques" on each
spacecraft in an array, from weak environmental effects, or from the required transverse
offset between the laser and the tether. Multi-element arrays don't need each
connection to be torque-free, as long as all the connections on each spacecraft null out
on the average. Periodic torques are easy to null out, with very small reaction wheels in
each spacecraft.
In the following sections, the details of the subsystems of the proposed formation flying
method are given.
B. Photon Thruster System
In this section the technical details of the intracavity photon thruster system shown in
Fig. 4 are presented.
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Satellite I Satellite II
Precision La ser
Power Meter
Laser Gai n Me dia
Ultrahigh Precision CW Photon Thrust
Intracavity Laser Beam
Len s
Diode
Pump
Laser
Pump
Lase r
Beam
Concave HR Mirrors
Figure 4. The schematic diagram of the photon thruster system.
The intracavity laser thruster system is proposed to be used to provide ultrahigh
precision repulsive force between satellites against the tether contraction force. If the
laser cavity is formed by two mirrors located separately in two satellites, the thrust, FT,
produced by a laser beam on each mirror is given by:
c
WRS
FT=, (1)
where W the is laser power, c the light velocity, 3 x 108 m/s, R the reflectance, and S is
the total power enhancement factor that is the ratio of the intracavity laser power to the
extracavity laser power. Here, the preferred laser cavity is a confocal resonator that
consists of two identical concave spherical mirrors separated by a distance equal to the
radius of curvature of the mirrors. The usage of the confocal resonator is much more
advantageous than that of a flat mirror resonator. The typical cavity with flat mirrors
requires an angular alignment adjustment accuracy of the order of one arc second.
However, the confocal resonator has a self-aligning property, thus the alignment
requirement requires only about a quarter of a degree, two orders of magnitude less
stringent than that with two plane mirrors. Furthermore, the former has much less
diffraction loss than the latter (Fowles, 1975).
The total laser power in the intracavity is a function of the reflectance of the HR mirror
and other complicated parameters, such as the saturation power of the laser media.
Here we consider first the effect of the HR mirror reflectance. Because laser photons
are virtually trapped in the intracavity laser formed between two mirrors, the average
laser power in the intracavity will be amplified. If there is no saturation of the gain media
and no thermal management limitations, the ideal total power enhancement factor, S, of
the intracavity is given by:
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2
)1(
)1(
RRT
S−
+
=, (2)
where R is the reflectance of the mirror, T is the transmittance through the mirror given
by 1 – R –A, and A is the absorption of the mirror coating during reflection. For high
quality mirrors, A ~ 10-6, thus, for the R < 0.99999, T ~ 1 – R, and the Equation (1)
becomes:
cR
W
FT)1(
2
−
≈. (3)
The parameters that determine the maximum attainable intracavity laser power are:
• The power saturation of the gain media
• The thermal management capacity of the gain media
• The HR mirror manufacturing consistency
For estimating the theoretical limit maximum intracavity laser power and the
corresponding thrust, the other parameters are neglected, and results of the maximum
theoretical thrusts as a function of the reflectance of the mirrors at the extracavity laser
power of 10 W are summarized in Table 1.
TABLE 1. The Maximum Theoretical Thrusts of the Photon Thruster as a Function of
the Mirror Reflectance at the Extracavity Laser Power of 10 W.
Maximum Operation
Laser Power
(extracavity)
HR Mirror Reflectance Maximum Theoretical
Thrust
10 W 0.90 - 0.99 (commonly used
in laser cavities)
0.67 - 6.7 µN
10 W 0.9997 (Newport
Supermirror)
220 µN
10 W 0.9999 (research grade) 0.67 mN
10 W 0.99995 (typically used
super mirror)
1.34 mN
The optimum design of the proposed intracavity photon thruster is different from that of
the typical laser cavities. The cavity design of the typical lasers is tailored to maximize
the laser output power in the extracavity. Depending on the characteristics of the gain
media, the reflectance of the output mirror (output coupler) is chosen 0.9 – 0.99 for the
conventional laser cavities. In some cases the HR mirror with the reflectance of 0.999
has been used (Lee, 2005). To minimize the absorption loss in the gain media, the
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proposed photon thruster should be designed to maximize the intracavity power, thus
the gain media should be very thin to minimize the absorption loss in the gain media,
similar to the one used in the state of the art solid state disk lasers used for intracavity
second harmonic generation, except without the need of the frequency doubling crystal
(Schielen, 2004). In this case, the thermal management of the gain media becomes an
important issue.
In this analysis, we have only considered the reflectivity and absorption loss of the
mirrors, however, several other factors including thermal limitation and optical
absorption and saturation of the laser gain media have to be considered. In reality,
because of the limitation in the laser gain medium and other thermal effect, the total
thrust presented in Table 1 should be considered as upper bounds. The current off-the-
shelf technological limit of the system reported to date is obtained with super mirrors
used for the cavity ring down spectroscopy (Romanini, 1997) (currently available in the
advanced research grade only) with the reflectance of 0.99995.
The maximum thrust of the proposed photon thruster as a function of the mirror
reflectance, R, is shown in Fig. 5. Note that the x-axis represents 1/(1-R), which is
approximately proportional to the number of reflections between two mirrors of the
photon thruster. The photon thrust shown here is calculated for a 10 W laser system,
and the higher laser power will reduce the required value for 1/(1-R) proportionally.
10 100 1000 10000 100000
0.1
1
10
100
1000
10000
Photon Thrust ( µN)
1
1 - R
Super Mirror
Predicted Capability of the Proposed System
()
FIGURE 5. The Maximum Thrust of the Proposed Photon Thruster of 10 W as a
Function of the Mirror Reflectance, R.
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Based on the currently available laser technology, by making the gain media thin
enough, the photon thruster with 0.999 -0.9999 is predicted to be readily possible with
the laser design optimized for maximizing the intracavity power in the near future. With
this, 10 W photon thrusters are predicted to be able to deliver up to 670 µN, which is
large enough to compensate various perturbations in the space environment for most of
missions envisioned as shown in Fig. 5. We note that the achievement of such high
photon thrust will require highly sophisticated gain medium and pumping design and
engineering, which is predicted to be within reach in the near future.
B.1. Photon Thruster Control Feedback Bandwidth
The proposed concept “inflates” and maintains the space structure using photons and
tethers. Let us compare the proposed system with the one formed by gas inflated
space structures that have been studied extensively and deployed in space. In the
former structure, the photons are trapped by mirrors and the structure is constrained by
tethers, while, in the latter structure membranes are used to both trap a gas and
constrain the structural dimensions. The typical bandwidth of the control feedback of
the structure depends on the medium that transmits the control. Because the photon
velocity is much larger than that of sound velocity in the gas by about 6 orders of
magnitude, the control feedback bandwidth of the former concept is much higher than
that of the latter. In the intracavity arrangement of the photon thruster, however, the
control feedback bandwidth can be much lower than that governed by the photon speed
alone, because the cavity acts as a capacitor. The bandwidth in this case is
proportional to the gain in the cavity. For example, with a gain of 10,000, the control
feedback band width of the photon thruster concept is larger than that of the inflated
membrane structure by only a factor of 100, which is still a very attractive figure. For
the system that requires faster control response, therefore, the gain of the cavity is
expected to be limited by the bandwidth factor.
B.2. Comparison of the Proposed Photon Thruster with other Microthrusters
We have investigated the detailed characteristics of the photon thrusters during this
period in comparison with those of other conventional microthrusters. The reason that
the photon thrusters have not been used is the thrust at a given propulsion energy is
much smaller than conventional thrusters. The ratio of the thrust, T, to the propulsion
power, P, is defined to be the specific thrust, TS:
P
T
TS=. (4)
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For nonrelativistic case, vmT •
= and 2
2
1vmP •
=, thus gIv
T
sp
S
22 == , where v is the
exhaust velocity and Isp is the specific impulse. For the photon case, c
P
T=, thus c
TS
1
=.
For both conventional thruster and the photon thruster, the specific thrust is inversely
proportional to the Isp or exhaust particle velocity; the higher Isp the less is the specific
thrust. Fig. 6 shows specific thrusts as functions of Isp of various conventional and
photon thrusters.
Isp (sec)
102103104105106107108
Specific Thrust (mN/W)
10-7
10-6
10-5
10-4
10-3
10-2
10-1
100
X 1,000
X 10,000
X 100
X 20,000
Electric Thrusters Photon Thrusters
Intracavity Multiplication Factors
Figure 6. Specific thrusts as functions of Isp of various conventional and photon
thrusters. The specific thrust of the photon thruster without the intracavity
momentum multiplication also follows the general trend of the most efficient electric
thrusters, and is several orders of magnitude smaller than that of the electric thrusters
with Isp less than 104. However, with the intracavity multiplication factors greater than
1,000, the specific thrust of the photon thruster rivals that of the most efficient electric
thrusters.
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The specific thrust of the photon thruster without the intracavity arrangement is 4 – 5
orders of magnitude lower than the conventional electric thrusters as shown in Fig. 1,
because of its high Isp. Therefore, the regular photon thruster is highly inefficient in
generating thrust, and this is the reason the photon thruster has been impractical in
most of missions. However, with the intracavity arrangement proposed here, the
momentum transfer, thus specific thrust, can be multiplied by “bouncing” photons
between high reflectance mirrors. If the number of the reflection of photons is greater
than 10,000, the specific thrust of the photon thruster becomes similar to the most
efficient electrical thrusters as shown in Fig. 6. Most importantly, the proposed
intracavity photon thrusters do not require rocket fuels, thus the mission ∆v is not limited
by the fuel capacity, but rather by the lifetime of the thruster, which is limited by the
pumping diode lifetime.
B.3. Lifetime of the Photon Thruster
When the interferometer in the satellite cluster operates, the intracavity laser should
operate continuously to dynamically adjust the relative distances and bearings between
the satellites. The lifetime of the formation system is thus limited by the lifetime of the
laser system. Currently, the lifetime of the diode pumped solid state lasers at full
operation power is limited by that of pump diodes to about 10,000 hours (1 year) for
continuous operation. With the reduced power operation or discontinuous operation the
lifetime of the pump diodes is expected to be longer. In any case, the overall lifetime of
the system can be further extended by simply replacing the pump diodes with new ones.
The alignment of the pump diodes does not have to be precise; a design with carousels
of pump diodes can be easily made. With a ten unit carousel, for example, the lifetime
of the system is extended to tens of years. Moreover, with the rapidly developing diode
laser technology, the lifetime is expected to increase significantly over the next decade.
Therefore, the lifetime of tens of years of the photon thruster is well within the reach of
the currently available off-the-shelf technology.
B.4. Maximum Range of Operation
The maximum range of the operation of the proposed system primarily limited by the
diffraction limit of the intracavity laser, thus the diameter of mirrors. The theoretical limit
of the intracavity length, L, for a confocal cavity resonator is given by (Yariv, 1975):
λ
21rr
L= , (5)
where r1 and r2 are the radii of the laser beam projected on the mirrors, and λ is the
wavelength of the laser. The required minimum diameter of mirrors for the photon
thrusters as a function of the operation distance is given in Fig. 7.
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Intersatellite Distance (km)
0.01 0.1 1 10 100
Mirror Diamter (cm)
1
10
Photon Thruster System: Mirror Diameter vs. Operation Distance
Figure 7. The required minimum diameter of mirrors for the photon thrusters as
a function of the operation distance.
For example, for MAXIM applications with L=200 m, the required minimum diameter of
the mirror is 3 cm. For the operation distances of 1 km and 10 km, the minimum
diameters of the mirrors are 7 and 20 cm, respectively.
The diameter of the mirrors that can be carried by microsatellites and equivalent
spacecrafts are probably limited by the weight of the mirrors. With the currently
available technologies, the weight condition sets the limit on the mirror diameter to tens
of cm. This in turn limits the maximum distance of operation to tens of km.
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B.5. Deployment and u-v Plane Activity Related Issues
The deployment process of the proposed system can be achieved by firing the photon
thruster at programmed thrust until the satellites establish a desired initial intersatellite
distance, while the pointing/alignment of the laser is actively controlled by the mirror
controlling system and the tether is gradually released. In the most of the envisioned
missions, the need for dense coverage of the u-v plane requires continuously variable
operation distance. In addition, in some missions, repeatable satellite segmentation
and desegmentation may be necessary.
So far, we have considered the design and performance of the proposed system at the
maximum intersatellite operation distance. Because the curvature and radius of the
mirrors of the proposed formation flying method are designed for the maximum baseline
distance, during deployment process and u-v plan activities at shorter distances, the
laser beam in the cavity will be defocused. In this case, the characteristics of the laser
cavity will be in between that of the confocal cavity and that of the flat mirror cavity.
The fractional power loss per transit of the laser beam in the intracavity is a function of
the Fresnel number, N given by (Fowles, 1975):
L
rr
N
λ
21
= (6)
where L is the length of the laser cavity, r1 and r2 are the radii of the laser beam
projected on the mirrors, and λ is the wavelength of the laser. In the proposed system,
r1, r2 and λ are constant, thus N is inversely proportional to L. The fractional power loss
per transit is a function of N, and for the confocal cavity, it is a rapidly exponentially
decreasing function of N, while for the flat mirror cavity, it is a slowly exponentially
decreasing function of N (Fowles, 1975). At shorter operation distances, as N increases,
the curve of the fractional power loss per transit shift from that of the confocal cavity to
that of the flat cavity. At very short operation distances, particularly during the initial
deployment, the laser cavity is close to that formed by flat mirrors. These effects of
increased N and the shift of curves on the fractional power loss per transit in the laser
cavity are expected to compensate each other; however, the degree of compensation is
not known currently. If the effect of the increased N on the fractional power loss per
transit under-compensates that of the curve shift, the increase of the mirror diameter is
necessary. The details of the optical analysis of these issues will be performed in
Phase II.
B.6. Photon Thruster Intracavity Mirror Alignment Issue
The laser system alignment at the onset of deployment is straightforward, because of
the proximity of the satellite will result in a large solid angle projected by the opposite
mirror. As the formation structure is inflated, the jittering of satellite positions may result
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in misalignment of the laser cavity. Therefore, the laser mirrors have to be actively
aligned as the cavity length increases during initial deployment stage. This active
alignment can be performed with piezo-crystals attached to the mirrors. The control of
the piezo-crystals will be achieved by active feedback signals from the photon detectors,
for example, four photo detectors covering four quadrants as shown in Fig. 8.
Satellite I Satellite II
Pre cisi on Lase r
Powe r Me ter
Laser Gain Media
Ultrahigh Precision CW Ph oton Thrust
Intracavity Laser Beam
Lens
Dio de
Pump
Laser
Pump
Laser
Beam
Concave HR Mirrors
Lenses
Piezo-Crystals
Laser Beam for
Interferometric Ranger
Photo Detectors
Mirror
Concave
HR Mirror
Figure 8. Schematic diagram of the HR laser mirror alignment system showing
the details of subsystems, including four quadrant photo detectors and lenses,
piezo-crystals, and a laser beam split for nano meter precision interferometric
ranging system described in the next section. The piezo-crystals align the HR
mirror based on the feedback signals from the four quadrant photo detector
outputs.
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However, after full deployment and during operation, if the alignment of the laser is lost
due to some major perturbations, including micrometeoroid impact, the reestablishment
of the laser cavity locking may be a non-trivial issue. In this case, the typical laboratory
procedure for establishing initial laser beams: rocking or scanning the mirror alignment
until the lasing starts. Once the onset of lasing is detected, the alignment mechanism
will tune or maximize the laser power and quality.
C. Interferometric Ranging System Integrated with Photon Thruster
System
In the conventional formation flying control, each spacecraft has to equip with both
lateral and longitudinal control systems including thrusters and metrology systems. The
use of crystalline structure in the formation flying simplifies the formation control, and
only longitudinal control and metrology systems are necessary. Furthermore, our
innovative design concept simplifies the system further by combining the propulsion
systems and metrology systems together.
The precision formation flying of satellites will require a highly sophisticated ranging
system for monitoring the intersatellite distance for the next generation NASA missions.
If the thrust of satellite maneuvering is provided by conventional microthrusters, the
ranging will require additional laser based interferometric system to the thruster control
system. The proposed formation flying method takes advantage of the dual usage of
the photon thruster as a laser source of the interferometric ranging system. This is
possible because in operation the intracavity laser of the photon thruster will be
operated with high stability ideal for the ultrahigh precision interferometric ranging
application. The proposed dual usage will significantly simplify the system design and
reduce the system weight and power consumption.
We have investigated the optimum way of combining the interferometric ranging system
with the photon thruster system. One of the best candidates for the ranging system for
the proposed ultrahigh precision formation flying method is the laser interferometric
ranging system. (Jeganathan, 2000, and Bender, 2003) Fig. 8 illustrates the schematic
diagram of the proposed subsystem with a Michelson interferometric scheme in which a
portion of the extracavity laser beam is reflected by a partially reflecting mirror (or a fully
reflecting mirror) to be used for interferometric ranging. The schematic diagram in Fig.
9 represents one of many possible designs, and the selection of the most suitable
system may depend on the specific mission requirement. The interference of the
primary laser beam in the primary satellite and the laser beam reflected by the mirrors in
the secondary satellite is used for assessing the relative distance change between the
satellites. This Michelson interferometry design of the ranging system is relatively
straightforward, but the ranging accuracy is in the order of tens of nms. The nm
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accuracy ranging will require more sophisticated interferometric ranging system that will
be discussed below.
Satellite I Satellite II
Precisio n Laser
Po wer Me te r
HR M irror HR Mirror
Las er G ain M edi a
Ultrahigh Precision CW Photon Thrust
Intracavity Laser Beam
Lens
Diode
Pump
Laser
Pump
La ser
Beam
HR Mirror
HR Mirror
Part al Mirro ri
Photodet ector
Part al Mirro ri
Part al Mirro ri
Figure 9. Schematic diagram of the photon thruster coupled with the laser
interferometric ranging system. Because of the dual usage of the photon thruster
laser for the interferometric laser, the system configuration and control becomes
simplified with reduced mass and cost.
During operation, the power of the photon thruster will be maintained ultrastable and the
force fluctuation due to the environmental perturbation will be countered by varying the
tension of the tethers with the precision translators. Upon encountering large
environmental perturbation or major realignment of the whole formation structure, the
tether vibration will be generated. The details of tether vibration countering scheme will
be extensively discussed in the later sections. The rapid countering of longitudinal
tether vibration can be achieved by modulating the laser power. Even in this case, the
interferometric ranging system can be designed to be insensitive to the overall laser
power fluctuations, because its operation relies on counting the interferometric fringes
rather than measuring the absolute power measuring. The more technical details of this
scheme are of Phase II topics.
The requirement for the absolute longitudinal metrology depends on missions, and will
not be considered in this report. In any case, it is highly technically challenge for the
design to achieve nm accuracy. Heterodyne interferometry is based on the production
of two coherent beams; a reference and a measurement beam with slightly different
frequencies. This can be done with two lasers or one laser with two acousto-optical
modulators (AOM). To simplify the architecture, we have chosen the latter design. The
schematic diagram of the proposed heterodyne interferometric metrology system in
combination with the photon thruster system is shown in Fig. 10.
In the proposed design, the first parts of these beams are sent to the reference detector
in which they interfere. The second part of the beams are sent to a reference
retroreflector and sent back to the measurement detector. The second par of the
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measurement beam is sent to the target retroreflector and is then sent back to the
measurement detector in which it interferes with the reference beam. The phase
difference between the signal on the two detectors, which is proportional to the path
difference, is transformed from the optical frequency region into an electrical one, and
the phase differences are measured. With the reference system located in the satellite
II and the measurement system located at the satellite I, using a return beam between
these satellites the relative distance variation can be measured to nm accuracy.
The present method represents one of many possible designs, and the selection of the
most suitable system may depend on the specific mission requirement. One of the
concerns is the possible interference between the two systems. During operation, the
power of the photon thruster will vary to counter the intersatellite distance change
resulted from the environmental force perturbation. However, the interferometric
ranging system is not sensitive to the overall laser power fluctuations, because its
operation relies on counting the interferometric fringes rather than measuring the
absolute power. Therefore, the possibility of interference between the two systems is
minimal.
Satellite I Satellite II
Precision Laser
Power M eter
HR Mirror HR Mi rror
Laser Gain Media
Ultrahigh P recision CW Photon Thrust
Intracavity Laser Beam
Lens
Diode
Pum p
Las er
Pump
Laser
Beam
Par t al Mir ro ri
Beam Splitter
Measurement
Det ect or
Ref erence
Detect or
AOM AOM
Retr ore flect or
Mirror
ODL
Figure 10. Schematic diagram of the photon thruster coupled with the improved
heterodyne laser interferometric ranging system capable of measuring the
distance to nm accuracy. Because of the dual usage of the photon thruster laser
for the interferometric laser, the system configuration and control becomes
simplified with reduced mass and cost. AOM: Acousto Optical Modulator. ODL:
Optical Delay Line.
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D. Tether System
D.1 Overview of the Tether System
The proposed intracavity laser system will be combined with a tether system that will
provide pulling-in force between a satellite pair through tension. For the gross length
adjustments of the tether, in one of the pair satellite is proposed to have a reel
mechanism and inchworm actuator that can be clamped to or released from the tether
to allow low-noise fine adjustments (Fig. 11).
Satellite I Satellite II
Tether
Clamp
Tether
Reel Inchworm
Piezo-Translator
Electromechnical Damper
Figure 11. The schematic diagram of the tether system.
The tether will be extended with the use of laser thrust that will be counterbalanced by
the tether tension. The other satellite in the pair is proposed to have a piezoelectric
translator with sub nm accuracy. Currently, off-the-shelf commercial piezoelectric
translator can deliver the accuracy resolution of 0.02 nm. Because the accuracy in the
distance maintenance relies on that of the piezoelectric translator, the proposed system
will be able to deliver the sub nm accuracy.
Let’s consider a 1-D system that consists of two spacecrafts connected with a tether. In
this case, the variation in length ∆lF with the tension FT, which is counterbalanced by
laser thrust, is given by:
l
A
F
Y
lT
F
1
=∆ , (7)
where Y is the Young’s modulus, A is the cross sectional area of the tether, and l is the
length of the tether. This can be rewritten as a Hooke’s law type equation:
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FT lkF
∆
=
(8)
where l
YA
k=.
If the tether is stretched by the photon thrust force, FL, the total force, F, applied on the
tether, thus spacecrafts, is given by:
for x>0,
PL FFkxF
+
−
=
(9)
and for x<0,
PL FFF
+
−
=
. (10)
Where V
lx ∆= and FP is the equilibrium force on the spacecrafts from the environmental
perturbation, such as gravitational gradient and solar pressure. In the normal operation,
by stretching the tether with the photon thruster force FL>Fp, the system operation will
be limited for the case x>0.
D.2. Nano-Meter Accuracy Baseline Control
FP can be countered by controlling either the photon thruster force, FL, thus the laser
power thrust, or the length of the tether. In the case in which FP is countered by
controlling FL, the tether length is kept constant by controlling the pump diode laser
power with the feedback distance signal from the laser ranging system. The finesse of
the distance accuracy in this case comes from the ability of fine tuning of the laser
power and stability. We consider a system that has 1 km baseline distance, thus a 1 km
Kevlar tether with a diameter of 4 mm. In this case, Y=1011 Pa, and k = 1.2 x 103 N/m.
For the change of the distance of 1 nm, the photon thruster accuracy should be 1.2 x
10-6 N = 1.2 µN. For example, for L2 orbit applications, FP is estimated to be less than
50 µN, and the average FL will be maintained around 100 µN, which will stretch the
tether by 90 nm. Therefore, the required thrust accuracy is 1.2 %, which is well within
the reach of the off-the-shelf laser power accuracy.
Another way to maintain the interspacecraft distance with 1 nm accuracy in the above
example is that the tether is stretched by the constant photon thrust of 100 µN in
average, and the perturbative force is counter balanced by changing the tether tension
by moving the piezoelectric translators. In this case, because the piezoelectric
translators have the resolution much better than 1 nm, the limiting step is the photon
thruster power accuracy, and to obtain 1-nm accuracy, the photon thrusters should have
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the noise to main power ratio in the order of 10-2. In the currently available CW laser
systems, the noise to main power ratio of 10-5 can be achieved. Theoretically, the
piezoelectric translator can be continuously controlled in real time by the feedback
distance signal from the laser interferometers that measure and monitor the intersatellite
distance continuously. Therefore, the proposed intracavity thruster system, in principle,
can provide the distance adjustment better than nanometer accuracy.
D.3. Tether Vibration Control
The proposed PTFF will be cable of providing near adiabatic CW position control of
spacecrafts, thus the photon thruster and tether control system will generate minimal
abrupt perturbation unlike the control system with conventional microthrusters that will
provide typically pulsed impulse bits. However, there are several other mechanisms
that may cause significant perturbations on the dynamics:
1) If a tether is highly loaded, and some filaments break, the result will be a "step
function" relaxation, followed by ringing.
2) A small micrometeoroid or debris strike to a tether will also break some strands,
and causes a step change plus ringing that will travel back and forth on the tether until it
is damped out.
3) Major reorientation of the whole formation structure.
Because the damping process will result in deadband situations and significant
consumption of the system power, one of the aspects of the PTFF dynamics is what can
excite and how one can damp these modes, in particular with the use of computer
simulation. Numerous excitation modes of tethers are anticipated in tethers connecting
both orbiting or stationary satellites/spacecraft in formation flying. These modes can be
classified into 4 major types:
1. Low-frequency tether bending modes
2. Medium-frequency "sprung mass" modes (ignoring the mass of the tether
itself)
3. High-frequency "taut string telephone" axial modes
4. Tethered end-mass attitude motion.
In orbiting tethers, the lowest frequencies are typically associated with
libration. Libration is likely to be of concern mostly not from the point of view of
dynamics but from the constraints that gravity gradient torques impose on the pointing
of interferometers and similar instruments. (That will require study, but it is distinct from
the dynamics issues that are being addressed here.) Low-amplitude libration has a
period of 0.577 orbit for in-plane swings and 0.5 orbit for out-of-plane dynamics. This is
~1 hour in LEO, or ~6 months in heliocentric orbit. Somewhat faster (typically 1-10
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minutes in LEO, depending on the tether length and the ratio of tether mass to payload
mass) are the transverse modes of the tether. Faster yet is the typical "fundamental
sprung-mass" mode, with the tether serving as a long, low-cross-sectional-area, low-
spring-rate, nearly-massless spring between two endmasses.
The highest frequencies associated with the tether itself are typically the string axial
modes, in which the tether acts like the taut string telephone that most children have
played with. Typical high-strength non-metallic tethers like Kevlar, Vectran, or Spectra
have sound speeds of 5-8 km/sec, so the lowest such frequency is typically a second or
less. Much higher frequencies are also possible. In the limiting case, nearly "square
wave" tension profiles can propagate along the tether, and reflect back and forth
between the ends. In a vacuum environment, damping is limited to that intrinsic to the
tether or provided by the end attachments or endmasses. If a tether is highly loaded,
and some filaments break, the result will be a "step function" relaxation, followed by
ringing. A small micrometeoroid or debris strike to a tether will also break some strands,
and causes a step change plus ringing that will travel back and forth on the tether until it
is damped out.
Efficient and realistic simulation of tether dynamics in space has been a daunting task.
The main idiosyncrasy of simulation of tethered systems is their "computational
stiffness." This refers to a radical difference between the highest and lowest relevant
frequencies in the system. This problem also occurs in other physical systems. One is
detailed simulation of electrons and ions in plasmas, due to the mass
difference. Another is simulation of the evolution of galaxies, due to the time-scale
difference of individual star motions vs the slow evolution of the massive core. Stiff
systems are computationally demanding, because you need short timesteps to do
justice to the high-frequency dynamics, but you need long simulations to see the effects
of the low-frequency modes.
Tether dynamics specialists have written a variety of simulation programs. Even the
same analyst will often write different programs to study different aspects of the
dynamics. Since the tether excitation frequencies are so far apart, the dynamics and
control for each of them can be studied somewhat independently to the first order
approximation. However, a systems perspective is needed to ensure that the "fix" for
one type of dynamics does not become a strong driver for disturbances of another
type. This will require a combination of analysis and simulation. This "integration" effort
might use a detailed existing simulation tool that is computationally too intensive for
most of the detailed studies. For example, Dave Lang's "GTOSS" (Generalized
Tethered Object Simulation System) may be appropriate for studying interactions
between modes. But note that GTOSS tends to run slower on a Cray than TAI's
"BeadSim" program does on a PC. (In cases studied at JSC, both programs gave
similar results.) BeadSim speeds up calculations by artificially damping the taut-string
telephone modes rather than keeping track of them. This is justified during SEDS
deployment, but not for many other cases. TAI has also written a "1-D" simulation
program to simulate the axial response of a long (even tapered) tether to capture or
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release of end-masses, with or without reeling. This model handles the taut-string
telephone modes that BeadSim cancels, and deals with damping in an explicit
manner. As with nearly all models, the model does include some artificial damping that
is intrinsic to the discretization part of modeling. But that artificial damping can be
quantified and compensated for by adjusting the explicit damping model.
There are three areas that will require particular care in this modeling effort. One is
modeling the weak environmental effects that can drive the dynamics. This includes
both "fast" effects like micrometeoroid impact, or slippage at the attachment points, and
also "slow effects" such as temperature changes with orientation. The issue is to
determine which of many very weak forces might be of most concern. This is likely to
require a combination of analytical and simulation work.
A second area is appropriately modeling the tether attachments, and any control
capabilities built into them. It seems likely that in many applications, one will want to
change the baseline length by significant amounts. This requires reeling, which will
probably be quite noisy (on the scale relevant to interferometry). It will be followed by
periods of "quieting the system down." We suspect that during non-reeling times, one
will want to actively clamp the tether, to prevent slippage on the reel or any drive pulleys,
and the resultant noise from such slippage. But such attachments are not likely to have
much damping. Hence one may want to add a piezo-electric or other actuator into the
clamp support, to provide agile short-stroke control to damp out the residual high-
frequency dynamics. We have been searching for the most suitable satellite simulation
program for testing the proposed concept. Currently, most of commercially available
simulation programs are not equipped with the tether dynamics.
The third area is about the appropriate division of labor between such a mechanical
actuator, and a photon thruster actuator. The piezo-electric actuator will take less
power and mass for a given force, but it will have a slower response than the laser
(especially at the far end). In addition, it cannot provide a separation force, as the
photon thruster can.
The proposed ultrahigh precision formation flying, therefore, needs the ability to damp
out tether vibrations generated by maneuvers and environmental perturbations. For
simplicity, in this report, we only consider two major types of tether vibrations that are
expected to be generated: longitudinal and transverse vibrations. More detailed tether
vibration studies will be undertaken in Phase II.
In general, longitudinal tether waves are readily damped by the tether material friction
and, if necessary, by modulating the laser power in the photon thruster. Transverse
vibrations that are typically excited by retargeting maneuvers can be damped out by
vibration dampers located near the attachment points of the tethers to the spacecraft.
We propose to use an electromagnetic damper incorporated into the tether clamp
system as shown in Fig. 9. In this conceptual design, the clamp acts as an impedance
coupling mass and the linear motion induced by the transverse tether vibration is
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coupled through an electromechanical shock-absorber-like damper. In this damping
mechanism, the linear motion stiffness and damping can be adjusted by changing
electrical parameters over a wide range.
The transfer of energy carried by the waves from the tether to the passive damper is
maximized when the tether wave impedance is matched to the damper impedance
Miller and Hall, 1991). The impedance of transverse tether waves, Z1, is given by
vZ
µ
=
1 (11)
where ν = (T/ µ)1/2 is the wave propagation velocity, T the tether tension, and µ the
tether linear density. The impedance, Z2, of the damper responding at the wave angular
frequency ω is given by
⎟
⎠
⎞
⎜
⎝
⎛−+=
ω
ω
k
midZ2 (12)
where m, d, and k are the mass of the damper, damping coefficient, and spring
constant, respectively. By equating Z1 and Z2 one obtains the results: (a) d=µ v, and (b)
ω = ω0 = (k/m)1/2.
The wave transmissibility function (Pain et al. 1983) has the same formulation of the
energy loss per damping cycle of waves propagating along a tether terminated with a
spring and dashpot massive damper, as computed in Beletsky and Levin (1993). Fig. 12
shows the fractional reduction of the wave amplitude per damping cycle (δA/A), derived
from the wave transmissibility function, for cases of maximum and minimum baseline
lengths of the formation flying structure (Lorenzini, Bombardelli, and Quadrelli (2001)).
Note also that the tangential velocity of the collector spacecraft (with respect to the
system center of mass) was assumed equal to 2 m/s and 0.05 m/s at the maximum and
minimum baselines, respectively, in order to maintain the angular rotation rate of the
interferometer approximately constant during an observation, as currently planned.
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Figure 12. Wave amplitude reduction per damping cycle vs. wave frequency for
different masses of the tether attachment point damper at the baseline length of 1
km and collector’s tangential velocity of 2 m/s. From Lorenzini, Bombardelli, and
Quadrelli (2001).
The result in Fig. 10 shows that if the damper is tuned to the first modal frequency of the
transverse waves, the first mode is damped in one damping cycle (Dcycle). However,
higher frequency modes take more than one cycle to damp out, but a damper with light
mass can damp very effectively higher-order modes over a few damping cycles,
especially at long baseline lengths. The analytical results were confirmed through
numerical dynamics simulations by Lorenzini, Bombardelli, and Quadrelli (2001) of the
interferometer at a 1-km baseline. Their result presented in Fig. 13 shows the
transverse oscillation amplitude, measured at the tether mid point after a quick (and
consequently perturbative) retargeting maneuver. The impedance-matched damping
system that is activated at t = 4,000 s rapidly abates the lateral oscillations as predicted.
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Figure 13. Tether transverse vibration damping from numerical simulations on the
mid-point tether amplitude. From Lorenzini, Bombardelli, and Quadrelli (2001).
Note that the vibration did not decay even 4,000 sec after the tether excitation.
However, the impedance-matched damping system that is activated at t = 4,000 s
rapidly abates the vibration.
D.4. Thermal Contraction/Expansion of the Tether System
One of important problems that the tether based system will encounter is thermal
expansion/contraction of the tethers due to exposure or non-exposure of sunlight. In
fact, this thermal expansion/contraction is the key factor that limits the usage of any
formation methods relying on solid monolithic structured beds, in which the engineering
of the real-time response system to the thermal effect is daunting. However, the
proposed system with tethers and photon thrusters has a built-in capability of
responding to such thermal perturbation. In the proposed tether system, the length
change ∆lt resulted from the temperature change ∆t is given by:
tllt
∆
=
∆
β
. (13)
For example, at l = 1 km an environmental temperature change, ∆t =10 C, will result in
the tether length change as much as 5.5 mm. Because the thermal perturbation will be
a very slow process (matter of hours or even days), it will be readily compensated by a
piezoelectric linear translator coupled to the tether in the similar fashion to the force
perturbation. The interferometer monitors the change and provides the feedback signal
that controls the piezoelectric translator for compensation. If the temperature change
results in the tether length changes greater than the dynamic range of the piezoelectric
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translator, a reel and/or stepper motor system will kick in to provide much larger
dynamic range.
D.5. Tether Lifetime Issue
The lifetime of the tether in the space environment is a critical limiting factor on the
lifetime of the mission with PTFF. For example, risk of tethers being cut by
micrometeoroids or orbiting man-made debris is an important issue for tethered systems
operating in LEO. Currently, experimental data on the rate of failure of a tether line is
limited to the results of the SEDS-2 experiment and the TiPS experiment in LEO.
Thus, it is not possible to assess the breakdown of the tether used in operations in other
orbits, such as MEO, LEO, and Lagrangian Orbit. In general, the lifetime in LEO can
be considered to be the lower bound for the lifetime in other orbits, thus, in this report
the breakdown of tethers in LEO operation is analyzed.
The micrometeoroid and debris risks vary differently with tether diameter. Debris is the
dominant risk for large tethers (>3 mm) at LEO altitudes above 400-500 km, on the
other hand, micrometeoroids are for thinner tethers or lower altitudes. There are wide
ranges in estimates of tether risk. One approach is based on SEDS-2 experience. The
19.7 km long SEDS-2 tether was cut 3.7 days after deployment, and the remaining 7.2
km length appeared to remain intact for the remaining 54 observable days of its orbit
life. This means there was 1 cut in 460 observable km-days of exposure of a 0.78 mm
diameter braided Spectra tether. Carroll and Oldson, [13] extrapolated the SEDS-2 data
to other sizes by scaling with crater size distribution data derived from LDEF and other
sources. The results fit the following simple expression:
Estimated MTBF = (Dt + 0.3)3, (14)
where the Mean Time Before Failure is in km-years and the tether diameter Dt is in mm.
This is the only available unbiased flight-based estimate of tether MTBF at present. The
actual MTBF could be much higher or lower. The above formula predicts that a 1 km
tether with a diameter of 1 mm has a lifetime 2.2 years, and a 4 mm tether 80 years.
The tether lifetime in other orbits is expected to be much longer, although there is not
enough technical data to estimate it accurately. Furthermore, the perturbation force is
expected to be much smaller, thus the system requirement should be much more
relaxed than that of the LEO operation. During Phase II, this aspect will be investigated
further.
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D.6. Nonlinearity of Tether Properties
Tethers have distinctly non-linear and non-ideal behavior at very low tension, in
particular in the mN range. A tether of order mm thick (thick enough to have a decent
probability of survival), or some kind of collection of narrower tethers with cross-links
(caduceus, Hoytether, or other), will both have distinctly non-Hooke's-law stress-strain
response at millinewton-level tensions. In this case, the effective modulus may be
orders of magnitude less than that at higher tension, because at low tension, most of
the strain associated with changes in that low tension is due to residual straightening of
the bulk tether and of the fibers that make it up. And that behavior will change with
temperature, time, and handling history.
The short-term predictability and hence the controllability proposed here may be
adequate, but modulus values may not be representative of lines that are under enough
tension to be nearly straight. It may be important during initial deployment (or after
unwinding tether to increase spacecraft separation in an array) to cause some sort of
brief tension spike that is long enough to get rid of the worst of the "curvature memory"
but without causing much acceleration of the spacecraft that would be hard for the
lasers to cancel out. Or it may be enough to pass the tether through a series of
"straightening rollers" after it comes off the storage spool, to minimize memory of
curvature. Overall, this aspect is of minor concern, and during Phase II, this issue will
be addressed in detail.
D.7. Ideal Tether for PTFF
The ideal tether for PTFF does not need much strength, but it needs environmental
resistance, including impact resistance to micrometeoroids. Spectra (or its European
equivalent, Dyneema) tethers they have the best impact resistance, however, one could
use nylon, if lower stiffness seems to work better, and if the nylon can hold up in
vacuum ultraviolet as well as polyethylene does. As far as tether constructions, there
are several property parameters that can be adjusted, which might have some
effect. .One parameter is choosing between hollow tubular braids and flat braids (with
several complete twists from end to end, so minor twist variations don't cause significant
changes in solar pressure). One could even go to a twisted-strand "thread-like"
construction, but I'm not sure that would be an improvement. If flat braids are used,
another adjustable parameter is the amount of twist put into the completed braid. We
don't want enough to significantly torque the endmasses, but we need some twist to null
out variations in solar pressure forces on the tether. Very low twist might make a
difference. A third parameter is the braid tightness. That would tend to affect the
amount of hysteresis in slip-stick effects, with tighter braids having more hysteresis
(which might be good). A fourth parameter is the selecting either Spectra vs Dyneema,
and what strand denier to use. The different materials and deniers have different fiber
diameters, and that might make a difference. A fifth parameter is the possibility of
adding either a lubricant (to ease slippage between fibers) or a small amount of some
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matrix (to try to PREVENT slippage). It's hard to get good lubricant properties over a
wide temperature range, and keep the lubricant on tethers exposed to hard
vacuum. And polyethylene tends to have low friction to start with, but it may be worth
testing tethers with Braycote (a very low vapor-pressure lubricant often used in
space). A matrix would make the whole construction somewhat springy in
bending. The last option is to make the tether out of a monolithic thin flat strip of
oriented polyethylene film. It might be like a thin strip from a material similar to an
ordinary grocery bag. It would be stiff due to the monolithic nature, but it could be thin
enough to be fairly flexible. During Phase II, we will investigate these issues further,
and recommend the optimum tether material and structure for PTFF.
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III. PTFF TECHNOLOGICAL READINESS ASSESSMENT
The present Phase I study estimated the NASA TRL readiness of the proposed concept
based on the open non-classified literature and information. The summary of the
assessment is provided in Table 2.
Table 2. The estimated NASA TRL for the key components of the proposed PTFF.
Necessary Technology TRL Assessment Note
Intracavity Photon Thrusters TRL 3 Analytical and experimental critical
function and/or characteristic proof-
of-concept
Interferometric Ranging
System
TRL 5 Component and/or breadboard
validation in relevant environment
Tether System TRL 6 System/subsystem model or
prototype demonstration in a
relevant environment (ground or
space)
System Integration and Control TRL 2 Technology concept and/or
application formulated
We have thoroughly studied the required technologies for implementing intracavity
photon thrusters. We have shown that the concept has been demonstrated in the
laboratory experiments and all the necessary components are readily available off-the-
shelf. The proof-of-concept has been demonstrated in such experiments. This puts the
TRL of the intracavity photon thrusters 3. During Phase II, the further engineering
studies will be performed to optimize the characteristics of the intracavity photon
thrusters for PTFF applications.
The nanometer accuracy interferomeric ranging system has been demonstrated in
breadboard setup in space-like environment. (Jeganathan, 2000, and Bender, 2003)
Therefore, this sets the TRL of the interferometric ranging system 5.
The tether subsystem has been demonstrated numerously in space, thus this sets the
TRL of the tether system 6.
The overall system integration and control of PTFF has been formulated, thus this sets
the TRL of the overall system integration and control 2. We are confident that this TRL
can advance fast, once the development program is sufficiently funded, because all
subsystems have higher TRL.
In terms of Research & Development Degree of Difficulty Scale the proposed concept
was classified under R&D3: II - III (moderate -high): Requires to optimize photon thrust
design based on the current laboratory system and system integration, and to develop
control system.
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IV. MISSION SPECIFIC APPLICATIONS
Station keeping maintaining the relative inter-spacecraft distance for the next generation
NASA missions to nm requires the knowledge on details of potential perturbative forces
in specific orbits. Two types of station keeping are important: 1) relative station keeping
and 2) absolute station keeping. The relative station keeping is for maintaining the inter
spacecraft distance to nm accuracy, which is handled by the proposed PTFF system. In
most of missions, the relative station keeping is much more important than the absolute
station keeping. However, this does not exclude the need of the absolute station
keeping, because it is needed for various major maneuvers, such as orbit correction
and slewing. For example, the frequency of the operational requirement for the
conventional orbit correcting thrusters depends on the mission nature.
For various reasons presented above, it is highly desirable to be able to perform the
absolute station keeping with the PTFF system without adding conventional thrusters.
According to our preliminary estimate, this is indeed highly feasible. For example, let us
consider 2-satellite 1-km-baseline formation flying system that would need 3 photon
thruster systems, and each photon thruster has a 100 W pump diode laser which
require 200 W input power. Let us assume that the weight of each satellite is 100 kg,
and the pump diode lasers are directly used for without the intracavity arrangement, for
slewing. The thrust force on each satellite due to the total 300 W photon emission of all
3 pump diodes is about 0.5 µN, and the corresponding acceleration a is 10-8 m/s2. If we
assume that the structure accelerate to reach the halfway point and then decelerate to
the final point, the time requires for slewing by is given by
a
L
t2
2= (15)
where L is the distance that each satellite to travel to have slewing. For 1 degree
slewing, L= 17.5 m, thus t = 1.18 x 105 sec, or 1.37 days. For 10 degree slewing, t =
4.33 days. Therefore, even with conservative power setting the major slewing can be
achieved in a reasonable time without the need of other conventional thrusters.
Although, the slewing is slow, the alignment accuracy with the PTFF is unprecedented.
The diode laser pump beam can be turn and off typically within 1 sec, or if more
precision is required, a mechanical chopper can be used to achieve 1 sec operation
time. In this case, the slewing angle accuracy would be 2x10-11 rad = 4 micro-arcsec.
The laser beam can be chopped to have 10 msec operation, the angle accuracy would
be 0.4 nano-arcsec. Therefore, PTFF is able to provide the unprecedented target
alignment accuracy. The scanning accuracy is limited by the 1 nm baseline accuracy,
and it is in the order of 0.1 micro-arcsec for 1 km baseline system
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The system requirement for orbit correction depends on the specificity of missions.
During Phase I, we only consider the relative station keeping requirement of formation
flying in various mission specific applications. The method of absolute station keeping
with the PTFF system will be studied more in detail in Phase II.
In this section, we show that the various perturbation forces on the relative position of
the formation can indeed be countered by PTFF system efficiently.
A. Low Earth Orbit (LEO) Applications
A.1. Atmospheric Drag Effect
In LEO, atmospheric drag is the dominant orbital perturbation, which removes energy
from the satellite and leads to a decrease in orbital altitude. The atmosphere drag, FD,
on a spacecraft in low earth orbit is given by the formula (LaPointe, 2001):
B
Vm
FD2
2
ρ
= (16)
where m is the spacecraft mass, ρ is the atmospheric mass density at a given orbital
radius, V is the orbital velocity, and B is the ballistic coefficient, given by:
ACm
B
D
= (17)
where A is the spacecraft cross sectional area in the direction of motion, and CD is an
empirical drag coefficient, with value typically ranging from 2 to 4.
The orbital velocity, V, is approximately given by:
R
GM
V=, (4)
where G is the universal gravitational constant, 6.6726x10-11 N-m
2
/kg
2
, M is the mass of
the earth, 5.976x1024kg, and R is the orbital radius with respect to the earth’s center.
By combining Eqns. 2 – 4, we obtain:
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RAGMC
FD
D2
ρ
= (18)
Fig. 14 plots typical drag forces encountered in LEO for a range of spacecraft cross
sections, at altitudes ranging from 200 km to 1000 km, which is an excerpt from the
reference (LaPointe, 2001). The microsat with a mass of 100 kg and a cross sectional
area of 1 m2, and total available power of 100 W, the final laser energy can be in the
order of 20 W assuming electrical power to laser photon power conversion efficiency of
20 %. Assuming the intracavity multiplication factor of 20,000 the maximum thrust
produced by the photon thruster is 2.68 mN. Thus, the photon thruster can counter the
inter spacecraft force up to 2.68 mN. Assuming, each spacecraft in formation may have
the differential drag up to the maximum drag, the present system can station keep in
relative motion down to about 400 km in LEO. Although the present system can keep
the absolute distance between the spacecraft with nm accuracy, it does not counter the
absolute drag applied to the whole formation flying assembly. Such an overall drag may
be occasionally corrected by conventional microthrusters.
Figure 14. Atmospheric drag forces encountered in LEO for a range of spacecraft
cross sections, at altitudes ranging from 200 km to 1000 km. Quoted from
LaPointe, 2001.
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A.2. Gravity Gradient Effect
More important than the atmospheric drag effect in LEO tethered formation flying
structures is the gravity gradient effect. Because the satellites in tethered formation
orbits at the same angular velocity, thus with same centrifugal force, the gravitational
force does not cancel out resulting in the gravity gradient effect. If we consider 1-D two
satellite configuration vertically aligned to the earth center of the mass, the gravity
gradient effect between the satellites, FGG is given by
2
0
3
ω
mLF TGG = (19)
where LT is the tether length, m is the mass of the satellite and ω0, the angular velocity
of the center of the mass of the formation is given by
3
0
0r
GM
=
ω
(20)
where G is the universal gravitational constant (6.673 x 10-11 Nm2/kg2), M is the mass of
the Earth (5.979 x 1024 kg), r0 is the radius of the system's center of gravity from the
center of the Earth (m) (Cosmo,1997). From this equation, one can obtain that for LEO
the equivalent acceleration per km of the tether length, aGG, is ~ 3.5 x 10-4 g/km, and for
GEO, ~1.6 x 10-6 g/km.
For 100 km satellites LEO formation with 1 km tether the maximum gravity gradient can
be ~35 mN, which is one order of magnitude larger than the atmospheric drag at the
altitude of 400 km. Therefore, in PTFF LEO applications, the dominant major
perturbation is the gravity gradient.
In the pentagonal pyramidal formation flying structure with a 1 km baseline distance, the
maximum perturbation experienced in a photon-tether subsystem is 35/5 = 7 mN. The
compensation of this force would require about 50 W photon thruster systems with the
intracavity multiplication factor of 20,000. This power requirement is still within the
power budget of 100 kg microsatellites, therefore, the PTFF can be used in LEO
applications, if the 20,000 intracavity multiplication is feasible. If the maximum
achievable intracavity multiplication factor is less, then the required power for the
photon thruster is larger.
B. Geosynchronous Earth Orbit (GEO) Applications
Although atmospheric drag forces are considerably decreased in GEO, other orbital
perturbations become more significant. These additional perturbations include effects
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on the satellite orbits due to the oblateness of the earth, earth triaxiality, sun-moon
perturbations, and, at sufficiently high orbits, radiation pressure from the sun.
However, the effect of solar pressure impinging on a satellite can induce the differential
force depending on the shape and orientation of the spacecraft. The solar perturbing
force, FS, on a satellite of mass M and surface area A is given by: (Lovell, 1973)
ASFS)1(
σ
+
=
(21)
where S is the solar constant at 1-AU, 4.5x10
-6
kg/(m-s
2
), and σ is the average
reflectivity of the satellite. Assuming a microsatellite mass of 100-kg, a cross-sectional
area of 1-m
2
, and an average reflectivity of 0.5 yields an in-pane maximum relative
perturbation due to solar radiation pressure of approximately 6.8x10
-6
N. This can be
easily countered the current system, which can provide the thrust up to 2.68 mN.
The maximum gravity gradient effect on PTFF with a 1 km baseline distance in GEO is
about 1.6 x 10-4 N, which is more than one order of magnitude larger than other
perturbation effects. Therefore, even at GEO, the dominant perturbation force is the
gravity gradient force. However, this can be countered readily with a 20 W photon
thruster system.
C. Lagrangian and Other Orbit Applications
For the future NASA astronomical observations, significant advantages in satellite
station-keeping may be achieved by positioning the formation-flying array in a stable
Lagrangian orbit or into a heliocentric earth-trailing orbit. A number of formation flying
concepts have been developed to take advantage of the stability provided by
Lagrangian orbits and heliocentric fall-away trajectories. For a planet moving around the
Sun, there are five Lagrangian points in space at which spacecraft remains in a stable
orbit with respect to the planet. In this case, because the spacecraft does not orbit the
planet, perturbation forces from atmospheric drag or planet triaxiality do not affect a
formation flying, and only solar radiation pressure should be countered. As shown
before, the relative perturbation due to solar radiation pressure can be readily countered
by the present system. Therefore, the proposed ultraprecision formation flying is also
suitable for Lagrangian and other orbit applications.
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V. EXEMPLARY MISSION STUDIES
In addition to redefining and simplifying the existing NASA mission concepts, such as
SPECS and MAXIM, the present concept enables other emerging revolutionary NASA
mission concepts, such as X-Ray Fourier Transform Spectrometer proposed by Dr.
Schnopper and New World Imager Freeway Mission proposed by Prof. Cash, which
searches for advanced civilization in exo-planets. As the present concept is more
publicized, many other exciting concepts are expected to follow. One of such possible
missions is the construction of ultralarge space telescope with diameters up to several
km for observing and monitoring space and earth-bound activities. In this Phase I
report we highlight several new potential NASA missions, and more detail studies on
engineering specificities of the existing NASA mission concepts and the following new
concepts will be performed during Phase II.
A. Ultralarge Adaptive Membrane Telescopes
Many of NASA’s future missions require to build much larger space telescopes than
those available today, such as the Hubble Space Telescope (HCT) with its 2.4 m
primary mirror. Although HCT has provided significant information about outer space
during the last 10 years, much higher resolution is required to unlock the secrets of the
universe. Because the diffraction-limited angular resolution of a telescope is
proportional to the aperture diameter, larger and larger size space telescopes are
desirable for both the astronomical and Earth-observation communities for the future
missions.
Currently, for a monolithic space telescope, the size of launch vehicles is the major
limiting factor, thus mass-optimal telescope construction is crucial in increasing the
telescope size. For a reflector telescope two major types of mass-optimal design are
currently researched in reducing the size and weight: 1) membrane telescopes and 2)
sparse aperture telescopes. The membrane space telescope would consist of a
reflective layer just thick enough to reflect the science wavelength. Such a thin mirror
will have little bending stiffness and will behave as a membrane over large diameters.
According to the physics of large telescope design for an on-axis, filled-aperture
reflector, the standard wavefront error RMS requirement is 1/10 the wavelength. For
the visible wavelength space observation, thus, the RMS position requirement across
the whole mirror is hundreds of nanometers. The structural support of the membrane
plays different interacting roles. First the rim must be positioned to tolerances similar to
the mirror requirements over the entire circumference. Next the rim support must
provide enough stiffness to be the reaction structure for the application of membrane
tension. Lastly, the rim support connects the membrane to the rest of the telescope
structure. Therefore the membrane mirror rim support must provide a highly accurate
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and stable mirror boundary condition while receiving disturbance inputs from other
locations in the spacecraft (de Blonk, 1999).
PTFF may be a key technology enabling an Adaptive Membrane usage for the future
ultralarge telescopes with the diameters of kms. Fig. 15 shows an artist’s concept of
such an ultralarge Adaptive Membrane telescope. The baseline polygon formation
structure (a pentagon in this example) of PTFF can create tension in the membrane
surface and actively change the curvature for a wide range of optical resolutions in
conjunction with the baseline structure rotation along the axis electric dipole interaction
in the membrane as in Stretched Membrane with Electrostatic Curvature (SMEC)
Mirrors. In SMEC, the curvature of the membrane mirror can be adjusted by the
electrostatic potential applied across the membrane. (Errico, 2002) A PTFF secondary
mirror or membrane can also be flown as part of the system. According to the research
data (Stamper, 2001), the useful area of the membrane with 5 point attachment is over
80 % of the total membrane area, thus the need of the continuous attachment is not
necessary.
Figure 15. An artist’s concept of such an ultralarge Adaptive Membrane telescope
formed by pentagonal pyramidal PTFF. The 5-point attachment membrane has
over 80 % optically useful area according to the recent research result. (Stamper,
2001) The small structure shown left of the PTFF telescope is a conceptual
design of NASA James Webb Space Telescope (JWST) for relative size
comparison.
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One of the major concerns about using the membrane for the telescope is the flatness
achievable with the currently available manufacturing technology. With the currently
available manufacturing technology, Polymeric materials with extremely low CTE such
as polybenzoxazole (PBO) have been predicted to have an onorbit surface in accuracy
of only 0.35 mm RMS, which is nearly 3 orders of magnitude larger than the required
RMS accuracy for visible wavelength space imaging. Once the membrane telescope
area is revitalized considerably in the near future, the RMS surface accuracy is
expected to become rapidly much smaller than the currently available one.
Even though the manufacturing surface RMS accuracy is achieved in the desired level
of several hundreds of nanometers, other environmental perturbation effects, such as
membrane vibration and thermal expansion/contraction will decrease the RMS surface
accuracy during the telescope operation. For example, membrane reflectors on the
order of 10 m diameter made from Kapton-E with ten times larger coefficient of thermal
expansion (CTE) than PBO, have been shown to have ~1 mm rms surface inaccuracy.
The surface RMS fluctuation of PBO is expected to be smaller than that of CTE. Exact
RMS surface accuracy fluctuation will depend on the mission characteristics and
membrane material. Holographic techniques have been improved considerably to the
point that surface inaccuracies on the order of 1 mm rms or less are perfectly
correctable to the diffraction limit. Hence, a spherical or parabolic mirror made of a thin
film of PBO in conjunction with the proper holographic correction can enable a 10m or
larger space-based aperture operating in the diffraction limit. (Palisoc, 2000)
Therefore, PTFF is a key technology in the creation of an Adaptive Membrane for
imaging, a system which can change the design of future telescope systems. Figure 14
shows a onceptual drawing of what this system may look like. A ring of PTFF spacecraft
can create tension in the membrane surface and actively change the curvature for a
wide range of optical resolutions. An PTFF secondary mirror or membrane can also be
flown as part of the system. PTFF is a critical technology because it has the ability to
uniquely change the shape, size, and tension of the membrane surface and possibly
surface components using electric dipole interaction.
B. New-World-Imager and Freeway Mission
The ultralarge PTFF membrane telescope can be used for many future NASA missions
for obtaining the pictures of planets around stars. For example, one of the most
ambitious future proposed NASA missions is the “Freeway Mission,” proposed by Prof.
Cash (Cash, 2005) that will be able to study the exo-planets in same way that LandSat
and other Earth-observing systems study the surface of the Earth. The exo-planets are
the earth-like planets formed in other star systems.
Science fiction authors have always assumed we would have to visit distant planets to
obtain images like Fig. 16, but they can, in principle, be captured from light years away.
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Such a telescope will of necessity be large, to collect enough light to resolve and
analyze small details on distant planets. To capture the needed signal takes square
kilometers of collecting area, that has considered to be a practical impossibility for the
foreseeable future to date. (Cash, 2005). The above mentioned PTFF ultralarge km
diameter membrane telescope should enable such a mission so that the mission cost
would be within the envelope of what NASA can afford.
Figure 16. From the movie Contact. Dr. Arroway gets a brief glimpse of this alien
landscape. Such images are possible from Earth, although, at the moment, very
expensive. However, PTFF is capable of providing the crucial step towards
obtaining such an image. Excerpt from Cash, 2005.
C. 1-D Formation Flying Structure for Fourier Transform X-Ray (FTXR)
Interferometer
Fourier Transform Spectrometry has revolutionized the optical spectroscopy in the last
century. In particular, Fourier Transform Infrared (FTIR) Spectrometry has been
extensively used for numerous scientific and engineering applications. The innovative
concept by Dr. Schnopper (Schnopper, 2006) to apply a Fourier transform X-ray
interferometer to the spectral diagnosis of hot, X-ray emitting, cosmic plasmas is a
revolutionary idea. This concept fits nicely into the goals of the Beyond Einstein long
range mission planning program. The feasibility of this concept relies on exploiting the
results from a suite of state-of-the-art technologies developed in four disparate areas:
nanopositioners with picometer precision; flat, ultra smooth (~0.05 nm rms) silicon
crystals; ultra thin (~1 µm) polymide membranes; and ultra thin (~0.8 nm) bi-layer
deposition technology. X-ray Fourier Transform Spectrometer concept for high
resolution X-ray spectroscopy “leap-frogs” the solid state, cryogenically cooled,
microcalorimeter spectrometers now under study for NASA’s CONSTELLATION-X and
ESA’s XEUS, advanced X-ray spectroscopy missions. We can achieve a spectral
resolution of one part in 104 (and set a goal of 105) from highly excited ions of C through
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Fe (with considerably higher energy resolution than a microcalorimeter). Instead of
merely seeking to identify constituents and their abundances in the hot plasma we will
be able to “resolve” their line shapes. Some of the science, thereby enabled, is based
on structural dynamics (velocity distributions) rather than on diagnostics. Fourier
Transform X-Ray Spectrometer will map the emission and absorption structure from
ions heavier than oxygen near black holes, in stars, and in the intergalactic medium. We
can make these high resolution observations on heavily absorbed objects, where the
resolution of wavelength dispersive spectrometers at high energies is low. We have
chosen to study the wavelength range (~0.1 to ~4 nm) that encompasses the principal
H- and He-like emission lines from abundant species from C through Fe. We require a
separation of >100 m between the ~1 m diameter X-ray telescope and the
interferometer. This requirement can be met by the development of an ultra-precision
boom or two satellites flying in tightly controlled formation.
The present Photon Tether Formation Flight (PTFF) method enables the most stringent
formation flying required by the above-mentioned Fourier Transform X-Ray
Spectrometer. An example of conceptual design of the overall 1-D PTFF system for this
mission is shown in Fig. 17. The sub-nm precision station keeping between two
spacecraft is performed by three photon thrusters, three tether systems, and three
interferometric ranging systems, which allows ultrahigh precision attitude and directional
control of spacecraft as well. The integration concept of PTFF and FTXR is illustrated in
Fig. 18. The PTFF enables the FTXR’s operation in space by providing sub-nanometer
accuracy one-dimensional formation flying structure with a baseline of ~10 km. The
more in-depth engineering details will be studied in Phase II.
Te t h e r
Interferometric Ranging Beam
SC I SC II
Photon Thruster Laser Beam
Figure 17. An artist’s concept of 1-D PTFF structured by two spacecrafts and
three photon thrusters, three tether systems, and three interferometric ranging
systems.
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Figure 18. The conceptual design of Fourier Transform X-Ray (FTXR)
Interferometer combined with Photon Tether Formation Flight (PTFF) method.
The PTFF enables the FTXR’s operation in space by providing sub-nanometer
accuracy one-dimensional formation flying structure with a baseline of ~10 km.
The figure only shows one photon thruster, interferometric ranging system, and
tether system, however, in the actual application, the three such combined
systems are required as shown in Fig. 17.
One of the technological concerns of 1-D structure is the bending modes or lateral shift
of the structure. Most of the envisioned mission concepts are insensitive to this bending
or lateral shifting. For example, the primary requirement for the FTXR is that the axis of
the telescope remains aimed at the interferometer with a precision of about 0.1 arcsec
(at 10 km) and 1 arcsec (at 1 km) (Schnopper, 2006). These tight specs will insure that
the beam will fall on the interferometer.
Let us assume that two spacecraft are separated by L, and there is a relative lateral
shifting of δy as shown in Fig. 19.
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L
L’
θ
yδ
Figure 19. The relationship between the increase in the baseline distance and the
lateral shifting of the 1-D formation structure.
In this case the angle θ is given by
L
y
δ
θ
~. (22)
The increase in the baseline distance, δL, is given by
L
L
LLL −=−
θ
δ
cos
'~ . (23)
At small angle θ, 1/cosθ is given by
2
1~
cos
12
θ
θ
+. (24)
Therefore, δL, is given by
LL 2
~
2
θ
δ
. (25)
For a 1 km system, L=103 m, θ =1 arcsec = 4.8x10-6 rad, thus, δL ~ 1.2 x 10-8 m= 12
nm. Therefore, if the baseline distance is maintained better than 1 nm, the overall
shifting angle can be controlled better than 0.1 arcsec for a 1 km baseline system.
For a 10 km system, L=104m, θ =0.1 arcsec = 4.8x10-7 rad, thus, δL~1.2x10-9 m= 1.2
nm. Therefore, the lateral shifting tolerance can be controlled by the present PTFF
system.
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VI. DESIGN OF PHOTON THRUSTER AND NANO-PRECISION
THRUSTER TEST STAND FOR PHASE II
The proof-of-the-principle of the photon thruster has been demonstrated in numerous
laboratory experiments by other researchers, who especially work on the cavity ring
down spectroscopy. In the cavity ring-down system, the laser pulses in the cavities are
routinely bounced between two high reflectance mirrors as much as 104 – 105 times
(Romanini, 1997). In other laser setups where the intracavity power multiplications are
used for frequency doubling and mixing. In this case, the intracavity laser power
multiplication factor is typically in the order of 100 (Lee, 2004). The photon thruster
operation range should be similar to that of the cavity ring-down laser system, except
the gain media is in the cavity as in the intracavity frequency doubling and mixing
system. To maximize the intracavity gain, the absorption through the gain medium and
reflective loss on the surface of the medium should be minimized.
One of the crucial parameters in engineering the photon thruster is the maximum
intracavity power multiplication factor, i.e., the maximum number of bouncing of photons
between the cavity mirrors. The typical present day laboratory and industrial lasers do
not operate in this cavity design parameter range, rather, they are designed to generate
the maximum extra-cavity laser beam power and quality. To maximize the intracavity
laser power, the gain medium should be designed much different from the existing laser
systems.
During Phase II, we plan to address the following issues regarding the intracavity laser
system engineering design:
1. The optimum laser system
2. The optimum design of the gain medium to minimize the absorption and
scattering loss through,
3. The optimum pumping system design (side pumping or end pumping),
4. The optimum design of the thermal management system of the medium
5. More detailed optical studies on the diffraction loss of the mirror,
6. More detailed optical studies on the optimum geometry of the HR mirrors.
Some aspects of these issues can be addressed theoretically with the use of existing
software, such as ZEMAX, however, the laser system and gain medium related issues
should be addressed experimentally. During Phase II, we plan to build a 1 W
extracavity power photon thruster that consists of an intracavity YAG laser pumped by a
diode laser to research the above issues. Initially, when we wrote the Phase I proposal,
we contemplated to investigate the photon thruster concept with a 10 W system for
Phase II, primarily to produce enough thrust to be measured the typical torsion thruster
stand.
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During this reporting period we have learned that this will cost most of the research
budget for Phase II leaving not enough resources for other crucial studies, such as
formation dynamics simulation studies. Since the importance of the realistic tether
simulation is crucial, we have to reduce the budget for the photon thruster
demonstration. In fact, for the optimum engineering parameter studies on the photon
thrusters in Phase II, a 1 W laser system would be sufficient. Currently, at a reasonable
cost, a diode pumped YAG laser with the maximum power of 1 W for the laser media
composed of a YAG, and diode laser pumping arrangement available commercially.
We plan to build the photon thruster demonstration prototype system with the off-the-
shelf Newport Supermirrors with a reflectance of 99.97 %. This photon thruster is
predicted to provide the thrust of 22 µN. To measure such thrust confidently requires a
thruster stand setup able to measure with tens of nN accuracy reliably. During Phase I,
we have designed such a thrust stand system that can be readily coupled with the
photon thruster for Phase II. The schematic diagram of the proposed propulsion system
is shown in Fig. 20. The thruster stand for measuring such thrust was very complicated
and expensive to build with the previous arts. However, Phipps and coworkers (Phipps,
2005) have just developed an ideal thruster stand for our proposed research. The
thruster stand is relatively straightforward to build and able to provide a 25 nN accuracy
with the use of laser interferometry.
Laser Power Meter
Conca ve HR Mirro r
HR Mirror
Laser Media
Intracavity Laser B eam
To rs io n
Fib er
Counter
Weight
V
acuum Chamber
Inte rfere nce Pattern
Low Power Laser
Corner Cube
Windows
Optical Fiber
Photo Detector
for Fringe Counting
Pump Laser Diode
Figure 20: 1 W experimental setup for demonstration of the preprototype photon
thruster capable of delivering thrust up to 134 µN in the nN thruster stand with 25
nN accuracy. The thrust is measured by a torsion fiber system coupled with laser
interferometer and corner cube in vacuum.
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In Phase II, the laser output will be measured with a laser power meter as a function of
the pump diode laser input power and photon thrust between two mirrors will be
measured with the nN thruster stand. In Phase II, we plan to start with a 50 cm long
intracavity, and try to develop the technique to increase the length of the intracavity to
several m. The intracavity photon thruster outside of the vacuum chamber will be
enclosed in an inert gas environment to minimize the intracavity absorption. The cavity
system is very sensitive to the vibration and structural noise of the system. An active
mirror controlling system will be developed to compensate the building and equipment
structure vibration. Such an active mirror controlling system will be highly crucial to the
real implementation of the system. In the real system, the mirror system perturbation
will result from the tether control system. In Phase II, the mirror control system will be
refined such that it can handle the vibration and noise from the tether control system
upon integration of the whole system.
Fig. 21 shows the vacuum thrust stand setup. The test stand setup will be similar to
that developed by Phipps et al. (Phipps, 2005). The entire setup will be mounted in a
vacuum test chamber. The details of the thruster test stand are shown schematically in
Fig. 20.
Concave
HR Mirror
Photon
Thruster
Beam
Flag
Interferometer
Laser Beam
Corner
Cube
Vac uum
Compatible
Oi
l
Counter
Weight
Torsion
Bar
Figure 20. Detailed schematic diagram of the thruster stand with 25 nN accuracy.
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The test stand will use a torsion bar, a critical damping attachment with a flag immersed
in diffusion pump oil, and an optical corner cube for interferometer. The measured
thrust F is given by
R
k
F
θ
= (26)
where K is torsional spring constant given by
L
Gd
k32
4
π
= (27)
where G is torsion modulus, L is effective length, J is the polar moment, and d is
diameter of the torsion fiber.
With the fused silica torsion fiber of 78 µm diameter, k = 194 pN-m/µrad and R =
0.155m, so that k/R = 1.25 nN/µrad. To have the thrust stand to have 25 nN precision,
the rotation sensor must resolve 20µrad bar rotation. An interferometer based on a
solid glass retroreflecting “corner cube” (described below) is the key to resolving rotation
of the bar. Critical damping is provided by a flag immersed in diffusion pump oil. The
retroreflector, a solid glass corner cube, with a 2.54cm diameter aperture will be used
for reflecting the laser beam. The two reflected beams from the front surface and from
internal reflection of the corner cube will be combined on a surface creating
interferometric fringes. A 5 mW, 532-nm near-diffraction-limited CW beam expanded to
15mm collimated diameter using a beam expansion telescope. As rotation occurs, these
fringes move radially outward or inward, depending on the direction of rotation of the bar.
Counting the passage of the fringes, which can be done visually or using the figure 3
setup, gives rotation. Phipps et al. (Phipps, 2005) reported that this setup was able to
resolve 20 µrad bar rotation resulting in the thruster measurement accuracy of 25 nN,
sufficient enough for the Phase II demonstration.
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VII. ROADMAP
A. Predictions and Limitations on PTFF
In science, whenever an instrument has increased its accuracy/capability by orders of
magnitude, it has opened up numerous applications, and often new scientific fields.
Some of the major revolutionary examples are listed below in the order of dimension:
1. Subatomic Dimension
High Energy Accelerators: By increasing the acceleration energy, thus decreasing
probing dimension in subatomic particles by orders of magnitude compared with
previous accelerators.
2. Atomic Dimension
Scanning Tunneling Microscopy/Atomic Force Microscopy: By increasing the scanning
resolution by orders of magnitude with minimal disturbing of the observed system
compared with other microscopy systems.
3. Molecular to Day-life Dimension
Laser: By increasing the spectral purity and the focusing capability by orders of
magnitude compared with incoherent light sources.
4. Astronomical Dimension
In the astronomical dimension, one of the most important goals is to increase the
imaging resolution of the target, which is inversely proportional to the diameter of the
observation apparatus aperture. The potential revolutionary breakthrough in the
astronomical dimension has been recognized by NASA to be the precision formation
flight that enables ultralarge aperture platforms that are able to increase the aperture
size by many orders or magnitude, surpassing the one constructible on earth at the
budget within reach. Size limitations on launch vehicle fairings leave formation flying
as the only option to assimilate coherent large apertures or large sample collection
areas in space (Leitner, 2004). In particular, the formation flight technology with the
ultrahigh baseline accuracy at 100 m to several km baseline lengths is the key
technology for enabling such ultralarge aperture platforms. Thus far, a solution for
maintaining a precise spacecraft configuration in space has proven illusive.
With the currently available microthruster technologies, the spacecraft formation
structure with the baseline accuracy in the spacecraft formation structure of 1 cm is
feasible. PTFF, if fully demonstrated, is predicted to increase the currently available
accuracy by seven orders of magnitude. It is too early to compare the potential of PTFF
in opening new concepts and possibly new scientific fields with the above mentioned
major revolutionary breakthroughs in 20th Century; however, PTFF is predicted to result
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in new concepts and scientific fields by quantum leaping in the baseline accuracy in the
spacecraft formation structure in the astronomical dimension.
By investigating the limitation of the technology, it is possible to predict the overall
potential of the proposed concept. In this section we review some of the limitations on
the eventual performance of PTFF in enabling new science and engineering concepts.
A.1. Limitation on the Baseline Size
The limitation on the PTFF baseline size is determined by the HR mirror diameter in the
photon thruster. With the currently available technologies, the HR mirror diameter in the
order of 1 m is possible. However, considering each spacecraft should carry possible 3
such HR mirrors, the weight of the mirrors probably sets the limitation, rather than their
size. For microsatellites, it is possible to carry three 20 cm diameter HR mirrors,
resulting in a baseline size limitation of 10 km. For a larger spacecraft missions, if the
weight of HR mirrors is not of major concern, the current mirror (1 m diameter)
technology sets, the baseline size limitation to be ~200 km. For the missions requiring
much larger baseline dimensions, such as LISA, PTFF is unsuitable.
A.2. Limitation on the Baseline Accuracy
The limitation on the baseline accuracy of PTFF is determined by the accuracy of the
tether control system and/or the laser interferometric ranging system. The accuracy of
off-the-shelf piezo-translators is in the order of 0.1 nm. The accuracy of the state-of-
the-art interferometric ranging system is in the order of 1 nm. Therefore, the current
limitation of PTFF baseline accuracy results from the limitation of the interferometric
ranging system, and it is in the order of 1 nm.
A.3. Limitation on the PTFF Lifetime
There are many factors limit the lifetime of PTFF. Two of the most important factors are
the lifetimes of the pump diode laser of the photon thruster and the tether. Currently,
the lifetime of the diode pumped solid state lasers at full operation power is limited by
that of pump diodes to about 10,000 hours (1 year) for continuous operation. With the
reduced power operation or discontinuous operation the lifetime of the pump diodes is
expected to be longer. In any case, the overall lifetime of the system can be further
extended by simply replacing the pump diodes with new ones. With a ten unit carousel,
for example, the lifetime of the system is extended to tens of years. Moreover, with the
rapidly developing diode laser technology, the lifetime is expected to increase
significantly over the next decade.
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The lifetime of the tether in the space environment is another critical limiting factor on
the lifetime of PTFF. In general, the lifetime in LEO can be considered to be the lower
bound for the lifetime in other orbits, thus, in this report the breakdown of tethers in LEO
operation is analyzed. The lifetime analysis of the tether based on the effect of
meteoroid impact predicts that a 1 km tether with a diameter of 1 mm has a lifetime 2.2
years, and a 4 mm tether 80 years. The tether lifetime in other orbits is expected to be
much longer.
In sum, with a reasonably thick tether (in the order of mms), the limitation factor on the
PTFF lifetime is mainly that of pump diodes. This can be readily increased to tens of
years with the use of a carousel design.
A.4. Limitation on the Number of Spacecraft
Theoretically or physically, there is no limitation on the number of spacecraft with PTFF.
For example, a 60 spacecraft structure with C60 fullerene structure (and other large
icosahedral structures) can be envisioned with PTFF. The limitation on the number of
spacecraft, currently, results from the economical affordability.
A.5. Limitation on the Orbit
PTFF relies on very subtle thrusts in the order of mN at the maximum with the total
spacecraft power budget of 100 W. This sets the limitation on the nature of orbits in
which PTFF can be used. One of the main limiting factors is the gravity gradient and
atmospheric drag, thus, higher orbits, such as GEO and Lagrangian Orbits are preferred.
If the mission requires LEO, PTFF may require higher total spacecraft power budgets or
the strategic coupling of PTFF with spinning the formation.
B. Required Technologies
Most of the technologies needed for implementing PTFF already exist and need only be
adapted. However, some technologies are more challenging than others. In this section
we list those technologies and discuss what needs to be done below.
B.1. Photon Thrusters
One of the most important technical issues related with photon thrusters is the
engineering of the intracavity system, in particular the gain medium system. The
required technology for this case is the optimum laser cavity design to obtain high
multiplication factors in the range of 1,000 to 10,000, and eventually over 10,000. On
the other hand, if the on-board power capacity of the spacecraft is more than 1W/kg,
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(with the use of nuclear power, such as RTGs, the power capacity can be much higher),
the multiplication factor can be much more relaxed. In this case, the required
multiplication factor is in the range of 100 – 1,000.
B.2. Nano-Meter Accuracy Interferometric Ranging System
The laboratory system for nano-meter accuracy interferometric ranging system was
demonstrated in a space-like vacuum chamber. The key technical challenge is in the
integration of the photon thruster laser with the design of the interferometric ranging
system. In particular, the laser beam off the photon thruster should have enough
stability.
B.3. Tether System
The tether system technology is more mature than other require technologies
mentioned above. The required technologies for the tether system are for pre-
tensioning to cure the tether, and the tether vibration controlling. Eventually, tethers
specifically designed for PTFF can be engineered.
B.4. Overall Dynamics Control System
The environmental perturbation from any direction will be distributed into the 1-D force
structure between paired spacecraft. 1-D structure analysis will be similar to 1-D
molecular vibrational analysis. For the 1-D system the lateral shifting or bending is a
major concern. However, we have shown here that the tolerance requirement to keep
the x-ray beam into the detector is much more relaxed, and 1 nm baseline (axial)
distance accuracy control can easily satisfy the lateral tolerance requirement for the
baseline length up to 10 km. The dynamics of this issue, however, need computer
simulation study. The 3-D dynamic simulation will require much more sophistication in
simulation.
B.5. Launch and Deployment
The system launch and initial deployment have to be investigated thoroughly.
Eventually, computer simulation would be ideal for this engineering issue. Launching of
the PTFF structure can be done in a single delivery or multiple deliveries. For multiple
deliveries case, the system should be assembled first before deployment. We have
shown that the photon thrusters can be used for providing the expulsion force required
for deployment. However, unrolling the tethers with minimal perturbation is an
engineering issue, and needs further studies.
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B.6. Target alignment, Scanning and Retargeting Issues
The method of scanning should be developed. For small angle scanning relative
changing of the baseline distances can create the scanning. The retargeting requires
major alignment angle of the formation, and requires additional thrust to the
interspacecraft thrust. It was shown preliminary in this report that the major retargeting
can be performed in a reasonable time scales (1 degree in a day) with the use of the
pump diode laser beams directly. Although, the slewing is slow, the alignment accuracy
with the PTFF is unprecedented with the accuracy of 4 micro-arcsec for 1 sec alignment
time. For 10 msec alignment time, the angular accuracy would be 0.4 nano-arcsec.
The scanning accuracy is limited by the 1 nm baseline accuracy, and it is in the order of
0.2 micro-arcsec. These are the order of magnitude estimate, and the more accurate
numbers will be provide with further mission specific studies.
C. Phase II and Beyond
Within the limitations of a Phase I investigation, we have tried to provide the theoretical
proof-of-concept, and identify and quantify the problems facing the realization of Photon
Tether Formation Flight (PTFF) for the next generation NASA missions. Although the
exposure of PTFF to public has been limited, it is already evident that PTFF enables
other new NASA mission concepts, exemplified by the X-Ray Fourier Transform
Spectrometer proposed by another NIAC fellow, Dr. Schnopper. In Phase II we
propose to further improve the confidence of PTFF that NASA can place PTFF in its
Roadmap.
A. Detailed Mission Design
During Phase I, we have produced a strawman design and first cut estimates of the
difficulties and problems to be faced in implementing PTFF in various exemplary
mission concepts. These difficulties and problems vary depending on the nature of the
missions. Logically, the next step is to create a detailed mission design utilizing experts
in all the various disciplines of space engineering to find an optimal solution to the
problem. During the initial stage of Phase II, we propose to generate a hierarchical list
of the future NASA mission concepts. Based on this list, we will try to create the
detailed mission designs for the top several mission concepts.
B. Development of Required Technologies
Several technologies are required to be demonstrated for the acceptance of PTFF
usage in a wide range of NASA applications. They include:
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1. the intracavity photon thruster,
2. the active HR mirror controlling system for the photon thruster,
3. the nano-meter accuracy interferometric ranging system.
4. the tether system with nano-meter control accuracy,
5. the dynamical control system that combines the photon thrusters and tethers
6. the tether vibrational control system
During Phase II we expect to further investigate these key technologies and decrease
the technical risk associated with each of these.
C. Implementation of PTFF in Actual Missions
The scientific and engineering community, like other communities, is often reluctant to
embrace a new idea and supplant older ones. During the peer review process for the
publication in STAIF in Phase I, we have answered numerous stringent technical
questions about PTFF, and have generated at least first acceptance of the concept in
the space community. Several reviewers scrutinized the concept, and at the end, they
agreed that there is no insurmountable technological blockage for PTFF so far.
The mission concepts that would first utilize PTFF are simplest ones, such as a1-D
PTFF structure for X-ray Fourier Transform Spectrometry in space, and a tetrahedron
PTFF structure that could be used for the modified structure of SPECS. Other relatively
simple systems are large membrane telescopes formed by PTFF. In particular, the
membrane telescopes with the diameter of 10 – 100 m will be within the reach in the
near future, once manufacturing issues related with the membrane uniformity is
resolved. We foresee that once this membrane telescope community is revitalized due
to enabling PTFF, the concurrent research on this aspect will speed up, and the desired
membrane quality will be within reach in the near future.
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APPENDIX
Publication in STAIF 2006 Proceedings
Space Technology and Applications International Forum, edited by M.S. El-Genk, AIP
Conf. Proc. AP813, pp.1213-1223 (2006)*
*Copyright (2006) American Institute of Physics. This article may be downloaded for personal use only.
Any other use requires prior permission of the author and the American Institute of Physics. along with
the following message: The following article appeared in (citation of published Article) and may be found
http://proceedings.aip.org/proceedings/confproceed/821.jsp.
A Contamination-Free Ultrahigh Precision Formation
Flying Method for Micro-, Nano-, and Pico-Satellites with
Nanometer Accuracy
Young K. Bae
Bae Institute, 1101 Bryan Ave., Suite C, Tustin, CA 92780, USA, www.baeinstitute.com
714-838-2881, ykbae@baeinstitute.com
Abstract. Formation flying of clusters of micro-, nano- and pico-satellites has been recognized to be more affordable,
robust and versatile than building a large monolithic satellite in implementing next generation space missions requiring
large apertures or large sample collection areas and sophisticated earth imaging/monitoring. We propose a propellant
free, thus contamination free, method that enables ultrahigh precision satellite formation flying with intersatellite
distance accuracy of nm (10-9 m) at maximum estimated distances in the order of tens of km. The method is based on
ultrahigh precision CW intracavity photon thrusters and tethers. The pushing-out force of the intracavity photon
thruster and the pulling-in force of the tether tension between satellites form the basic force structure to stabilize
crystalline-like structures of satellites and/or spacecrafts with a relative distance accuracy better than nm. The thrust of
the photons can be amplified by up to tens of thousand times by bouncing them between two mirrors located separately
on pairing satellites. For example, a 10 W photon thruster, suitable for micro-satellite applications, is theoretically
capable of providing thrusts up to mN, and its weight and power consumption are estimated to be several kgs and tens
of W, respectively. The dual usage of photon thruster as a precision laser source for the interferometric ranging system
further simplifies the system architecture and minimizes the weight and power consumption. The present method does
not require propellant, thus provides significant propulsion system mass savings, and is free from propellant exhaust
contamination, ideal for missions that require large apertures composed of highly sensitive sensors. The system can be
readily scaled down for the nano- and pico-satellite applications.
Keywords: formation flying, photon thruster, intracavity, satellite, spacecraft, interferometer, interferometric
ranging, propellantless, tether, SPECS, MAXIM, TPF, ST-3, contamination free, micro, nano, pico.
PACS: 07.87.+v, 95.55.-n, 95.55.Fw, 93.85.+q.
INTRODUCTION
In recent years, microsatellites and nanosatellites provide an opportunity to insert sophisticated sensors and
processing technologies into orbits of interest at low costs (Leitner, 2004). Building a cluster of small satellites has
been recognized to be more affordable, robust and versatile than building a large monolithic satellite. Specifically,
the grouped satellite cluster is crucial for enabling orders-of-magnitude improvements in resolution and coverage
achievable from advanced remote sensing platforms. Size limitations on launch vehicle fairings leave formation
flying as the only option to assimilate coherent large apertures or large sample collection areas in space (Leitner,
2004). For example, for NASA applications, the ultrahigh precision satellite clusters can be used for interferometry
and distributed large aperture sensors, especially at optical (TPF, and SPECS) and x-ray wavelengths (MAXIM)
(Leisawitz, 2004, Cash, 2002). For non-NASA applications, the proposed system can be used for advanced
geophysical monitoring where GPS and standard laser range finders are currently inadequate to measure and
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Photon Tether Formation Flying (PTFF) Bae Institute
monitor small changes in the movement of earthquake plates, and gravity wave detection. Other commercial and
military applications include distributed large aperture optical and infrared sensors for ultrahigh resolution
monitoring and imaging at low-cost.
Such a technology critically depends on the formation flying method that enables precision spacecraft formation
keeping from coarse requirements (relative position control of any two spacecraft to less than 1 cm, and relative
bearing of 1 arcmin over target range of separations from a few meters to tens of kilometers) to fine requirements
(nanometer relative position control). For example, one of the most challenging applications for formation flying
thus so far is that of the proposed x-ray interferometry for space imaging applications, MAXIM (Cash, 2000). The
concept has evolved to include a pathfinder mission, consisting of a single x-ray interferometer and a trailing
imaging satellite, and the full MAXIM, consisting of a fleet of 33 x-ray mirror satellites, a trailing collector satellite,
and an imaging or detector spacecraft. Summary of the requirement of the baseline accuracy tolerance of several
exemplary missions compared with the capability of the present formation flying method is shown in Fig. 1.
10-9 10-8 10-7 10-6 10-5 10-4 10-3 10-2 10-1 Meter
MAXI
M
SPECS TP
F
Present FF Metho
d
FIGURE 1. Required Base Line Accuracy of Several Exemplary Missions and the Capability of the Present FF Method.
In MAXIM, the relative distance between the hub satellite and collector satellites should be precisely maintained
with the tolerance of a few nm (10-9m) at the distance of 200 m, and the precision requirement in maintaining the
distance, thus, is 10 parts per trillion, one of the most stringent accuracy requirement seen in any scientific fields. In
addition, potential contamination of neighboring spacecraft by propellant exhaust plumes and the possibility of
pulsed electromagnetic interference with low power inter-satellite communications remain a real concern for
grouped satellite clusters. These requirements essentially rule out the usage of the most of the conventional
propellant based propulsion systems, such as gas hydrazine thrusters, pulsed plasma thrusters, hall thrusters,
electrostatic ion engines, and field emission electron propulsion systems.
To alleviate these concerns, several propellant-free formation flying methods have been proposed. The propulsive
conducting tethers and spin-stabilized tether systems have been proposed in place of on-board propulsion systems to
form and maintain satellite formations (Johnson, 1998, Quinn, 2000).
While such concepts offer intriguing
possibilities for small arrays consisting of only a few spacecraft, i