ARISE (Advanced Radio Interferometry between Space and Earth)

ARISE (Advanced Radio Interferometry between Space and Earth)

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ARISE (Advanced Radio Interferometry between Space and Earth)
http://www.lgarde.com/assets/content/files/publications/arise.pdf
JPL Document # 16330
1
Document prepared by the ARISE Team:
Art B. Chmielewski - Preproject Manager
Jim Ulvestad (NRAO) - Preproject Scientist
Pradeep Bhandari - Cryocoolers
Robert Chave - Subreflector design
Bob Freeland, Paul Willis - Antenna materials
Todd Gaier - Science instruments
Henry Garrett - Space environments
Rick Helms - Structures, inflatable antenna cognizant engineer
Mike Jones - Ground systems
Carol Lewis, Sal DiStefano, Gene Wester - Power
Leo Lichodziejewski (L’Garde) - Inflatable structure
Roger Linfield, Rick Wietfeldt - VLBI and VLBI processing
Bob Miyake - Spacecraft thermal design
David Murphy - Orbits and coverage
Muriel Noca - Systems
Yahya Rahmat-Samii, Robert Hoferer (UCLA) - Antenna design and RF compensation
Vincent Randolph - Avionics
Larry Roe (Univ. of Arkansas) - Inflation system
Sam Sirlin, Marco Quadrelli - Antenna dynamics, Attitude control
Dan Thunnissen - Propulsion
Charles Wang - Telecom
Space VLBI consultants:
Robert Preston
Joel Smith
The purpose of this document is to summarize the technical work performed by
the ARISE Team in FY’98. The work focused on the space segment of ARISE.
Ground segment will be studied in more detail in FY’99 and will be available in
the next edition of this document. Roadmaps for all the technologies involved in
the ARISE mission can be found in separate documents.
JPL Document # 16330
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TABLE OF CONTENTS
1. ARISE Mission Description p. 4
2. ARISE Science p. 5
3. Mission Design, Coverage and Constraints p. 8
3.1 Nominal orbit and sensitivity
3.2 Precession of the orbital elements
3.3 Space environment
3.4 Orbit trade-offs
3.5 Launch capability and sequence
3.6 Space environment
4. ARISE Inflatable Antenna p.16
4.1 Antenna general description
4.2 Reflector configuration
4.3 ARISE structures and thermal analyses
4.4 Antenna surface precision
4.5 Inflation system
4.6 Deployment sequence and canister design
4.7 Subreflector description
5. Science Payload p.27
5.1 Science data requirements
5.2 Receivers/Amplifiers
5.3 RF adaptive compensation
5.4 System performance
6. Spacecraft Description p.40
6.1 Spacecraft configuration
6.2 System description, mass and power budgets
6.3 Gain and observation duration budget
6.4 Spacecraft data flow
6.4.1 Avionics
6.4.2 Telecommunications
6.5 Spacecraft thermal design
6.5.1 Cryocoolers stage
6.5.2 Bus thermal design
6.6 Spacecraft attitude control
6.7 Structure and mechanisms
6.8 Power subsystem
6.9 Propulsion subsystem
7 Ground systems and mission operations p. 66
8. Cost p. 68
JPL Document # 16330
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Appendices p.72
A: ARISE mass budget
B: ARISE power budget
C: ARISE structures and thermal analysis
D: Solar Electric Propulsion system
E: ARISE radiation environment
F: ARISE electrostatic discharge (ESD) environment
G: Power Subsystem
H: ARISE cost estimates
I: ARISE study team
JPL Document # 16330
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1. ARISE Mission Description
40,000km
Space
Antenna Earth
Telescopes
(NRAO, DSN, etc.)
13-hr period
ARISE (Advanced Radio Interferometry between Space and Earth) is a space Very Long Baseline
Interferometry (VLBI) mission consisting of one (or possibly two) 25-meter radio telescope(s) in a
high elliptic Earth orbit. In conjunction with arrays of ground telescopes, ARISE will image the
most energetic astronomical phenomena in the universe, namely supermassive black holes. The
mission objectives are to image radio sources with a resolution of 10-20 microarcseconds, which
corresponds to an improvement in resolution over today’s Space VLBI mission by two orders of
magnitude. ARISE’s observing bands will be 8, 22, 43, (60), and 86 GHz, and system noise
temperatures down to 10-20 K. Science data will be downlinked at a rate of 1-8 Gbps
The ARISE spacecraft is placed in a high elliptical Earth orbit in order to synthesize the largest
possible imaging aperture. The nominal orbit has a perigee altitude of 5000 km and an apogee
altitude of 40000 km. The mission lifetime is approximately 3 years, with a potential start in 2005
and launch in 2008. The spacecraft launch mass is about 1700 kg, allowing for a launch to the
desired orbit with a Delta II class launch vehicle.
The ARISE spacecraft is designed with two primary goals: 1) to make the mission as low cost as
possible, 2) to maximize the antenna performance. Several innovative ideas were used to make the
mass and volume a minimum. An inflatable antenna is baselined with a mechanical antenna as
secondary option. An inflatable antenna can be packed into a volume that is about 100 times less
than an equivalent mechanical structure; the inflatable also is about 5 times lighter and 6 times less
expensive to manufacture. The inflation system was combined with the attitude control and
propulsion system for additional mass savings. All the structural elements of the antenna will be
rigidized after the deployment, to nullify the need for any supplemental gas. The reflector structure
also will be operated at a pressure of 1/10,000 atmospheres to lower the requirement for make-up
gas, which will be needed to replace the gas lost by leakage due to micrometeoroid penetrations.
The inflatable antenna peformance is enhanced by using mechanically shaped secondary reflector
and an adaptive feed.
To further lower the cost of the mission, ARISE will take advantage of technology development by
other missions and programs. It will use the coolers and low noise amplifiers which are being
developed for the Europe-led Planck mission, scheduled for launch in 2007. The data systems will
take advantage of the developments by the ground VLBI and DOD programs. The inflatable
antenna technology development will be greatly aided by the Space Inflatables Program.
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2. ARISE Science
The primary goal of ARISE is the study of the environment of black holes and other compact
objects, as well as the disks of matter surrounding these objects. Secondary goals are the studies
of gravitational lenses throughout the universe, and of coronae in active stellar systems.
Massive black holes are believed to be the power sources for active galactic nuclei, including the
gamma-ray “blazars” first detected by the Compton Gamma Ray Observatory. Among the
questions of scientific interest are the method of feeding these black holes, and how they use the
fuel to generate the light-speed jets seen in blazars. ARISE will image the region of primary
energy deposition and delivery in these objects with a resolution of light days to light months,
depending on the blazar distances. Observations at 43 and 86 GHz are required to image these
regions in optically thin emission, so that our view is not restricted to an opaque surface. Imaging
of these regions in polarized radiation will map out the inner magnetic field structures, required for
understanding the energy-generation processes. The combination of the ARISE imaging with
gamma-ray observations and X-ray spectroscopy is particularly important to provide a complete
picture of the highly energetic phenomena near massive black holes.
Fig. 1.1
An important corollary to the study of black holes is the study of accretion disks on a variety of
spatial scales (Fig. 1.1). Such disks are the reservoirs of fuel for black holes, other compact
objects, and star-forming regions. Understanding the physics of these disks, and the relation
among disks of various sizes, is critical to understanding the complete life cycle of matter near
massive objects. ARISE will image the 22-GHz water megamaser emission at the centers of active
galaxies, providing direct measurements of black hole masses and of the physics of the accretion
process (Fig. 1.2). Weaker maser emission from disks in galactic star-formation regions will also
be imaged to help show how the accretion phenomenon scales with mass and power of the
accreting objects. Finally, continuum radio imaging of superluminal jets associated with energetic
x-ray binary stars in our galaxy will probe the accretion processes on much smaller scales than is
possible in extragalactic objects.
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Fig. 1.2
A key secondary science goal for ARISE is the study of gravitational lenses throughout the
universe. The accessible resolution of tens of microarcseconds (Fig. 1.3) provides sensitivity to
compact objects in the mass range of 104
to 106
solar masses; no other astronomical technique has
access to such objects, which are among the candidates for the “missing” dark matter in the
universe. These lenses also can be used as “cosmic telescopes”, since their magnification provides
enhanced linear and angular resolution of distant objects. Thus, it may be possible for ARISE to
have an effective resolution even better than that indicated just by the size of its orbit and its
observing frequencies.
Another important secondary goal for ARISE is the study of the coronae of active star systems.
Very sensitive imaging at 5 GHz or 8 GHz will allow mapping of these coronae with an effective
linear resolution much smaller than a stellar radius. In addition, motions in coronal mass ejection
events more powerful than those on our Sun can be followed in time scales of hours. Of particular
interest is the capability of imaging stellar flares to search for brightness temperatures that indicate
coherent emission processes in the coronal plasma; such high brightness temperatures can only be
accessed using VLBI baselines much larger than an Earth diameter.
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Fig. 1.3
ARISE Science Advisory Team
Members from U.S. Institutions:
Prof. Moshe Elitzur, University of Kentucky
Dr. Lincoln Greenhill, Harvard-Smithsonian Astrophysical Observatory
Prof. Jacqueline Hewitt, Massachusetts Institute of Technology
Dr. Arieh Konigl, University of Chicago
Dr. Julian Krolik, Johns Hopkins University
Dr. Roger Linfield, Jet Propulsion Laboratory
Prof. Alan Marsher (Chairman), Boston University
Prof. Robert Mutel, University of Iowa
Dr. Susan Neff, NASA Goddard Space Flight Center
Dr. Robert Preston, Jet Propulsion Laboratory
Dr. Jonathan Romney, National Radio Astronomy Observatory
Dr. Ann Wehrle, California Institute of Technology
Members from Foreign Institutions:
Dr. Denise Gabuzda, Lebedev Physical Institute, Russia
Dr. Michael Garrett, Joint Institute for VLBI in Europe
Dr. Leonid Gurvits, Joint Institute for VLBI in Europe
Prof. Hisashi Hirabayashi, Institute for Space and Astronautical Science, Japan
Prof. Russell Taylor, University of Calgary, Canada
Prof. Esko Valtaoja, Tuorla Observatory, Finland
JPL Document # 16330
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3. Mission design, coverage and constraints
3.1 Nominal orbit and sensitivity
The ARISE orbit is one of the major factors in determining the science return from the ARISE
mission. The final orbit selection will be made after a detailed trade-off between scientific goals and
spacecraft design, and after the VSOP and RadioAstron results. In the meantime, a nominal orbit
for ARISE has been specified by the science requirements, along with a range of possible values
for each parameter. Table 3.1 summarizes these parameters.
Quantity Nominal Possible Range
Semi-major axis 29,000 km 15,000 - 50,000 km
Eccentricity 0.6 0.25 - 0.75
Apogee Altitude 40,000 km 40,000 - 100,000 km
Perigee Altitude 5,000 km 1,000 - 6,000 km
Inclination 60 deg 30 - 63.4 deg
Orbital Period 13.5 hr 5 - 30+ hr
Perigee Precession 6 deg/yr 0 - 280 deg/yr
Node Precession 21 deg/yr 5 - 180 deg/yr
Orbit Knowledge 10 cm 3 - 20 cm
Table 3.1: Orbit parameters for ARISE
The selection of the perigee altitude should allow for overlap between ground-ground and ground
space telescope baselines for calibration purposes. Low perigee (near 1000 km) will cause more
rapid precession of the orbit plane. The apogee altitude should be high enough to provide
information at the desired resolution, and will result from a trade between high angular resolution
and high dynamic range imaging.
There are several different approaches to examining the ARISE sensitivity. The approach we adopt
here is a quasi-physical approach. Given the 7-σ sensitivities from ARISE to a single VLBA
antenna of 1.7, 4.3, 13.8, and 110 mJy at 8, 22, 43, and 86 GHz respectively, we may ask the
question on a given ARISE baseline, what type of sources are we able to detect. In this analysis, a
source is represented as a single Gaussian component of total flux density, So, and brightness
temperature, To. In Figure 3.1, the detection limits for the 4 ARISE observing bands for 3 different
values of baseline length (20,000, 40,000, and 80,000 km) are shown. The straight line
corresponds to when the visibility function reaches a value of 0.9. Thus, to the right of this line
even though a source might be detected, we would be unable to determine its size. With this simple
Gaussian source model, the minimum brightness temperature that can be detected on a baseline of
length D with a detection limit of Sd is given by:
Tbmin = 3.1 x 108
(D/104
km)2
(Sd/mJy) K
This minimum detectable brightness temperature is for a source with a flux density of 2.7Sd. From
Figure 3.1, we can see which sources can be detected as a function of observing frequency and
baseline length. As an alternative to representing a Gaussian source by its total flux density and
brightness temperature, we can represent it by its total flux density and FWHM size. In Figure 3.2,
we show what sources can be detected as a function of these two source parameters and baseline
length. From this figure we can see the size scales that will be probed by the different ARISE
observing bands. In this figure the 0.9 visibility straight line is different for each observing band.
JPL Document # 16330
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Fig. 3.1: Detection limits for the 4 ARISE observing bands as a function of source flux density (So)
and maximum brightness temperature (Tb) for 3 different values of baseline length. Sources at the
right of the curved lines are detectable, while sources to the left of the diagonal lines are resolvable.
Fig. 3.2: Detection limits for the 4 ARISE observing bands as a function of source flux density (So)
and FWHM size for 3 different values of baseline length. Sources above the curved lines are
detectable, while sources below the diagonal lines are resolvable.
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3.2 Orbit normal and (u,v)-coverage
One of the prime goals of the ARISE mission is to image sources with unprecedented angular
resolution. The highest angular resolution (u,v)-coverage is obtained for sources that lie along the
orbit normal and anti-normal directions. In equatorial coordinates (α, δ) these directions are given
by (Ω - 6h, 90o
-i) and (Ω + 6h, i-90o
). In Figure 3.3, we show the (u, v)-coverage obtained for a
one orbit observation with ARISE in its nominal orbit and the VLBA as a functional of the
equatorial co-ordinates of the source. Note, that the (u,v)-coverage is essentially linear when the
source lies in the orbit plane, shown as a sinusoidal curve in Figure 3.3. Due to the nodal
precession (dΩ/dt) the position in the sky of the orbit normal and anti-normal directions precess
with time. In the nominal orbit for ARISE, this precession period is 15.7 years compared to 1.61
years for the current Japanese space VLBI satellite, HALCA. For the nominal mission lifetime of 3
years, Ω only precesses by 70o
, implying that some directions of the sky would never have
outstanding (u,v) coverage. The science tradeoffs involved in such a situation are under discussion
(also, see section 3.4).
Fig. 3.3: All-sky (u,v)-coverages for a one orbit observation with ARISE in its nominal orbit and
the VLBA.
3.3 Precession of the orbital elements
At the moment, both the injection argument of perigee ωo and the right ascension of the ascending
node Ωo are free parameters. However, for the assessment of the ARISE orbit environment, a
value of ωo = 0o
has been assumed. The nominal orbit has an orbital period (T) of 13.56 hours and
the precession rates of ωo and Ωo are +6o
/yr and -23o
/yr respectively for 60o
inclination (and
+63o
/yr and -40o
/yr respectively for 30o
inclination). These nominal ARISE orbit precession rates
can be compared to the HALCA precession rates for both ω and Ω which are +353o
/yr and -228o
/yr
respectively. The relatively low precession rate for ω may not be a problem provided that there are
no spacecraft link constraints except for the requirement that ARISE must be above the elevation
limit of a tracking station. However, the injection value of ω, ωo needs to be further studied as the
optimum value depends on the geographical distribution of tracking stations. If we assume DSN
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tracking (with 2 tracking stations in the northern hemisphere and one in the southern hemisphere)
then ωo = 180 is to be preferred over ωo = 0 for orbits with i < 63.4o
and hence dω/dt > 0. The low
precession rate of Ω is of some concern and will be further addressed in Section 3.4.
3.4 Orbit trade-offs
It is instructive to examine how the derived orbital parameters P, dΩ/dt, and dω/dt depend on hp,
ha, and i. In Figure 3.4, we show how the orbital period (P) depends on hp and ha. The nominal
orbit has a period of 13.56 hours, which coincides quite nicely with the typical ground-based
VLBI observation length. Increasing the orbital period much beyond this value has some
disadvantages, since the typical imaging observation must last at least one orbit, and radio source
structures may vary on time scales appreciably shorter than 24 hours. In Figures 3.5 and 3.6 we
show the nodal precession rate dΩ/dt for orbit inclinations of 60o
and 30o
respectively. By
lowering the inclination from the nominal 60o
to 30o
we increase the nodal precession rate, for a
given hp and ha, by a factor of sqrt (3) (Ωdot is proportional to cos i). However, this in itself, only
reduces the nodal precession period from 15.7 years to 9.06 years. Reducing hp from 5,000 km to
1,000 km (while keeping ha at 40,000 km) and reducing i from 60o
to 30o
reduces the nodal
precession period to 3.95 years, which is comparable to the mission lifetime.
One consequence of lowering both the perigee height and the inclination is to increase the perigee
precession rate (since dω/dt is proportional to 5cos2
i - 1). At an inclination of 60o
, dω/dt is very
low since this inclination is close to i = 63.4o
, where dω/dt = 0. For hp = 1,000 km, ha = 40,000
km, and i = 30o
, ωdot = 145o
/yr. With these orbital parameters over the 3 year ARISE mission
lifetime there are 2.48 ω precession periods. With DSN tracking, ARISE will be able to be tracked
longer when ω = 270o
compared to ω = 90o
. Thus, in this case, an injection value of ωo = 180o
would to be preferred over a value of ωo = 0o
.
Fig. 3.4: Orbital period as a function of perigee and apogee heights.
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Fig. 3.5: Nodal precession rate dΩ/dt as a function of perigee and apogee heights for i=60o
.
Fig. 3.6: Nodal precession rate dΩ/dt as a function of perigee and apogee heights for i=30o
.
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Fig. 3.7: Perigee precession rate dω/dt as a function of perigee and apogee heights for i=60o
.
Fig. 3.8: Perigee precession rate dω/dt as a function of perigee and apogee heights for i=30o
.
In conclusion, nodal precession rate can be increased significantly by lowering the perigee height
from 5,000 to 1,000 km and lowering the inclination from 60 to 30o
.
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3.5 Launch capability and sequence
The launch vehicle selected for the ARISE mission is the McDonnell Douglas Delta 7925. The
7925 version features 9 solid rocket motors, and a Star 48B spinning third stage. Its delivery
capability can be summarized as:
- 1720 kg on a 185 x 40000 km altitude orbit, i=28.7 deg., 3-m dia. fairing
- 1330 kg on a 185 x 40000 km altitude, i=90 deg., 3-m dia. fairing
- 1220 kg on a Molniya orbit (370 x 40000 km), i=63.4 deg., 2.9-m dia. fairing
- 1170 kg on a Molniya orbit (370 x 40000 km), i=63.4 deg., 3-m dia. fairing.
Figure 3.9 shows the injected mass as a function of apogee altitude and fairing type for the 3-stage
7925 vehicle. For ARISE, the 2.9-m diameter (9.5-ft) fairing was selected since it allowed enough
space for the stowed spacecraft and since the injected mass was larger than the 3-m, leaving more
margin for spacecraft mass growth. Figure 3.10 shows the performance capability of the 2-m dia.
fairing as a function of inclination. Ultimately, the choice of the inclination for the ARISE orbit will
depend on the spacecraft mass.
Once launched, the spacecraft will go through a de-spin and stabilization mode. A perigee raise
maneuver will then occur at the GTO (Geo Transfer Orbit) apogee. Then several sequences of
deployment will happen: deployment of the inflatable antenna and solar arrays; deployment of a
rigid astro-mast type arm to carry the sub-reflector to about 3.6 m from the spacecraft; and finally
deployment of the telecom antenna.
Fig. 3.9: Delta 7925 three stage launch vehicle capability
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Fig. 3.10: Injected mass as a function of inclination. Delta 7925 2.9-m fairing.
3.6 Space environment
The radiation environment for two different inclinations and two different arguments of perigee
was assessed. In summary, the trapped magnetospheric charged particles (electrons and protons)
dose behind 100-mils of aluminum was:
- 105 krad[Si]/year @ i = 30 deg., ω = 0o
.
- 40 krad[Si]/year @ i = 60 deg., ω = 0o
.
- 102 krad[Si]/year @ i = 30 deg., ω = 90o
.
- 60 krad[Si]/year @ i = 60 deg., ω = 90o
.
The solar flares proton dose have been calculated to be on average about 10 krad[Si] for 3 years,
which is small compared to the trapped magnetospheric particles. The requirement for the reflector
radiation material resistance was assessed. Surface dose on the reflector was estimated at about 130
Mrad/year, while bulk dose adds up to about 40 Mrad/year. More details on the radiation
environment can be found in Appendix E.
The electrical surface charging (ESD) of the ARISE main reflector was also investigated. In
summary, differential static potential between the two thin Kapton sheets (one Al coated) forming
the main reflector could reach about 20 kV under worst conditions, which could lead to selfsustained
arcs. Although the use of an ITO coated Kapton sheet for the canopy might be
satisfactory, electrostatic discharge still remains a materials issue until appropriate tests are done.
Details on the ESD environment can be found in Appendix F.
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4. ARISE Inflatable Antenna
An inflatable antenna was chosen for the ARISE main reflector because of mass, cost and low
storage volume considerations. However, a lower risk approach is currently being evaluated using
mesh antenna types. Results of this investigation will be reported in the second edition of this
document.
4.1 Antenna General Description
Fig 4.1: ARISE antenna depicting innovative technologies
The ARISE primary reflector is comprised of a reflective membrane with an RF transparent front
canopy to complete the inflated lenticular envelope. The lenticular is combined with a tubular
peripheral support torus forming a large, lightweight, yet relatively stiff space structure. To
minimize membrane stress, the lenticular structure is pressurized with < 4 x 10-4 psi of N2 and
attached to the torus ring with constant force springs in a “trampoline” fashion.
Figure 4.1 shows the configuration of deployable/inflatable/rigidizable technologies used by the
ARISE antenna to meet the large structure, low cost, low mass, low storage volume performance
requirements.
The antenna assembly is aligned and attached to the spacecraft using three tubular support struts
that are inflation deployed from a stowage canister located at the top of the S/C bus. At the torus
the antenna support struts are kinematically attached at 120° intervals. They are designed with
Low mass inflatable
support structure
Low mass inflatable
solar array
Adaptive feed
Low mass RF subreflector
Common gas generation
system (Propulsion, ACS,
Inflation)
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optimum diameters, wall thickness, and lengths to give mechanical rigidity (bending, torsion) yet
minimal obscuration and shadowing of the primary (see Fig. 4.2).
Recent advances in thin film photo-voltaics and the demonstration of the L’Garde ITSAT Inflatable
Solar Array have been incorporated into the ARISE design. To meet its power needs, ARISE
requires a large solar array, but at the same time needs low array mass and stowed volume to meet
launch vehicle constraints. The ITSAT Solar Array with its high packaging efficiency, low areal
density, and solar blanket/inflatable frame design is the best available technology to meet the
ARISE requirements.
Rigidizable support structure technologies are used wherever possible on ARISE in order to
minimize the need for “make-up” inflation gas. The primary reflector torus and antenna support
struts maintain their mechanical stiffness and shape by using cold rigidizable rubber and/or
polymer materials that solidify (phase change) in cold space environments. To maintain low
temperatures and minimize thermal gradients these support members will be wrapped in MLI
blankets. Because of high temperature conditions, the ARISE solar array blankets will be
supported by thin walled polymer struts laminated with aluminum foil. The strut tubes are initially
over-inflated past the aluminum foil’s yield point. Once the pressure is removed, the stressed
aluminum maintains much of the strut’s rigidity and shape.
4.2 Reflector Configuration
The current reflector configuration was selected after extensive evaluation of various reflector
antenna geometries, and resulted from some early trade-offs between on-axis and off-axis
configurations. Figure 4.3 pictures the two different configurations and Table 4.1 summarizes the
pros and cons. Upon thorough examination, it was determined that an off-axis configuration
offered better science performance (less obscuration) and fewer constraints for the rest of the
spacecraft design.
Side Spt. Struts (Cold Rigid.)
23.1 m x 20”∅ x 12 mil wall
Bottom Spt. Strut (Cold Rigid.)
8.2 m x 14”∅ x 12 mil wall
Torus (Cold Rigid.)
10.5”∅ x 12 mil wall
Primary Reflector (Pressurized)
28.4 m x 25 m x 0.5 mil/membrane
Solar Array Spt. Struts (Al Rigid.)
6.9 m x 5”∅ x 5 mil wall
Solar (Hi Eff Si Cell) Array Blanket
6.9 m x 2.6 m (ala ITSAT)
Subreflector Spt. (Gr/Ep)
3.9 m
Subreflector (Gl/Cyanate)
1.65 m, e = 0.555 S/C Buss (Al)
3.5 m x 2.5 m ∅ (in structural model)
Fig. 4.2: ARISE structural configuration diagram
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Fig. 4.3: On-axis versus off-axis configurations
On-axis Pros Off-axis pros
FOV No Obscuration
Structures Symmetry
ACS Ease of control
RF Un-Obscured Pointing
Fabricability Same complexity and reflector RMS Same complexity and reflector RMS
Table 4.1: On-axis versus Off-axis trade-off
Four different off-axis configurations were then evaluated (Prime Focus, Gregorian, Cassegrain,
and Schmidt Cassegrain (see Figure 4.4)) using the following criteria: dynamics/structural
stiffness, thermal stiffness, RF performance, mass, complexity, deployment reliability and
alignment. It was determined that the Gregorian off-axis design uses a smaller and less complex
structure and secondary reflector than Cassegrain types. Furthermore, the Gregorian off-axis
system offers the possibility of a mechanically shaped secondary that is reasonably sized and
controlled.
Prime Focus Gregorian Cassegrain
Fig. 4.4: Various off-axis configurations
Based on these arguments, a Gregorian dual-reflector antenna system was selected for ARISE.
Figure 4.5 displays a vertical cross-section through the reflector configuration. All the dimensions
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shown in this figure are in meters. The geometrical parameters, which fully define the reflector
configuration as shown in Figure 4.5, are given below:
On-axis “mother” reflector diameter D = 50 m
On-axis “mother” focal length F = 11.55 m
Off-axis sub-aperture diameter D = 25 m
Tilt angle between main reflector and subreflector axis β = 5.67 deg.
Inter foci distance L = 2.4 m
Subreflector eccentricity ε = 0.555
Figure 4.6 shows a three-dimensional representation of the reflector configuration. A ray, coming
in from the antenna bore-sight direction, reflected off the center of the main reflector and the
subreflector and received by the antenna feed at focus, is also sketched.
Z(main)
X(main)
25.000
12.500
11.550
2.388
1.538
Figure 4.5 (left): Vertical cross-section through the dual reflector geometry as proposed for
ARISE.
Figure 4.6 (right): Three-dimensional representation of the ARISE dual reflector geometry.
Additionally a ray incoming from the bore-sight direction, reflected off the main and subreflector
and received at the focus is sketched.
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4.3 ARISE Structures and Thermal Analyses
A major consideration in the determination of the relative merits of the off-axis Prime Focus and
the Gregorian configurations was structural and thermal behavior. Of special concern were
dynamic response, inertial static loading, and thermal distortion effects on antenna shape and
alignment. Because of the “soft” nature of inflatable structures, and the temperature sensitivity of
polymeric membranes, Structural and Thermal analytical models were created that had more
resolution than simple static diagrams and lumped masses. The structure model consisted of over
700 elements (Appendix C) that closely approximated off-axis parabolic curvatures, strut
orientation, subreflector alignment, solar array geometry, and S/C mass distribution. Special
attention was given to membrane elements, polymer properties, and tubular geometries; all critical
to inflatable structural behavior. The thermal model was based on the same nodal geometry and
properties allowing a direct one-to-one correspondence between temperature profiles and structural
elements.
Analytical Results : When the Prime Focus and Gregorian performance predictions were
compared, there was very little difference across the board. The Gregorian antenna dynamic
performance fared slightly better than the Prime Focus since its center-of-mass was closer to the
S/C bus and its antenna support struts were shorter. Inertial response (static thrust) and thermal
distortions were virtually the same between the two configurations. Consequently, the Gregorian
was selected over the Prime Focus for other reasons than structural and thermal performance
(obscuration, corrective optics, …).
ARISE’s dynamic response was analyzed using normal modes analysis (force driven vibration
stimuli have not yet been specified). Table 4.2 lists the first six non-rigid body normal modes for
the off-axis Gregorian. In general, the modal frequencies were shown to be representative of large
space structures, and are within acceptable bounds given ARISE’s operating scenarios. The
resultant modal shapes are classic with tip displacements that are non-critical. (Refer to Appendix C
for coordinate references.)
Mode # Modal frequency
(Hz)
Modal Shape Max. Tip
Displacement (cm)
1 0.3 Primary & S/C bus
“nodding” to each other
12.5
2 0.5 Primary Y tilt 13.5
3 0.8 Subreflector Y
cantilever
5.5
4 0.9 Subreflector Y tilt 5.8
5 1.0 Primary “trampoline”
motion
6.1
6 1.2 Subreflector X tilt 2.5
Table 4.2: ARISE Normal Modes Analysis
Figure 4.7 shows the resulting structural displacements due to a conventional, static (without
transients) thrust maneuver loading. A thrust vector of 0.015g at 2° off-axis was applied at the
base of the S/C bus in order to study worst case asymmetric inertial loading. A maximum of 4 mm
displacement was predicted. Since thrust maneuvers will not be performed during science
observations, it was felt, based on these preliminary results, that thrust/slewing maneuvers would
not create critical/castastrophic stress or strain conditions.
JPL Document # 16330
21
Off-Axis Gregorian
Thrust Analysis: Displacement
Torus Constrained in
Translation
0.015 g Thrust,
gymbolled 2° off-axis
meters
Fig. 4.7: Structural displacements due to a static thrust maneuver loading
After the initial nominal thermal analysis showed no difference between the Gregorian and Prime
Focus configurations, a more detailed worse case orbital thermal analysis of the off-axis Gregorian
structure was performed. One of the sub-solar points of the ARISE elliptical orbit was chosen for
this test case. It was felt that the combination of solar heating at the bottom edge of the lenticular
antenna structure with the Earth’s albedo/IR would generate large gradients across the reflector
membrane. The results shown in Appendix C, Frame 5 indicate a large gradient of approximately
115 C°. The resultant thermal distortions from this thermal profile are currently being analyzed.
4.4 Antenna surface precision
An essential element of the ARISE design is the reflector precision. L’Garde has built a 7 meter
reflector with a 1.7 mm RMS accuracy but is predicting 1mm RMS accuracy on reflectors of 25m
or more with appropriate development effort. To generate a credible estimate of the magnitude and
shape of the 25m ARISE antenna error, ground measurements from the 14m Inflatable Antenna
Experiment (IAE) reflector were utilized. The IAE reflector shape is shown in the top right of
Figure 4.8. It was measured during a ground test in preparation for flight. During the test, one of
the torus supports slipped and was not discovered until after the measurements were taken.
Nonetheless it is considered representative of the types of errors seen in this class of reflector albeit
somewhat exaggerated. As a projection of the type of errors that will be seen in future reflectors
which are expected to be more systematic, a FAIM software model prediction of the reflector shape
was utilized. FAIM predicts the shape of the inflated reflector analytically, but includes no random
errors as seen in material properties and manufacturing errors. The two reflectors were interpolated
together resulting in the reflector shape representing the convolution of the theoretical global error
generated by FAIM and the manufacturing errors scaled from the as-measured IAE.
gymbaled
JPL Document # 16330
22
FAIM Result
0.7mm RMS
IAE Measurement
Errors “Softened to 0.7 mm RMS
Interpolated ARISE Baseline (1mm RMS)
•FAIM prediction of systematic
error or “W shape”
•Contains no manufacturing errors
or material inconsistencies.
•Errors “softened” for combination
with measured surface.
•IAE measurements contained a flaw due to poorly
placed torus supports during testing.
• “Random” errors are representative but the magnitude
is too high
•Errors were “softened” to represent current
manufacturing and design techniques and proper torus
attachment.
•ARISE baseline represents latest design and
manufacturing techniques
•Includes measured errors from IAE.
•Includes systematic error or “W shape”
•Expected RMS error of 1mm is a prediction
of future accuracies based on current work
and future development efforts.
DZ mm
Figure 4.8: ARISE Reflector Precision Projection
4.5 Inflation system
The inflation system for the ARISE mission must provide gas for initial inflation of the struts,
torus, envelope (reflector/canopy assembly), and solar array booms, and make-up gas for the
envelope over the life of the mission. A range of options was considered, including tanked gas,
chemical gas generation, and combinations of these. The baseline specifications for the inflatable
antenna are 3.0 kg necessary for initial inflation plus the appropriate amount of make-up gas
required over a three-year mission life, and an envelope operating pressure of 10E-4 psia.
The system as currently configured utilizes tanked gas for initial inflation and catalytic hydrazine
decomposition for make-up gas. The tank masses are based on scaling relations from a prior study
(Thunnissen 1995). No redundancy is incorporated into this conceptual design. The gas tank is
0.24m-dia, T-1000 aluminum-lined graphite-epoxy, initially contains 0.4 kg He gas at 6000 psia,
and has a mass of 1.4 kg. The hydrazine tank is 0.40 m-dia titanium, initially contains 23 kg LHZ
at 500 psia, and has a mass of 2.4 kg. Estimated mass of catalyst, valves, regulators, filters,
orifice, and associated plumbing is 1.2 kg, yielding a total wet mass of 28 kg. With the plumbing
items located between the two tanks, the system will fit into an envelope of approximately 0.11 m3
.
The components are essentially off-the-shelf items, although some development of the catalyst bed
is anticipated to minimize the ammonia content of the products, to provide the minimum molecular
weight possible. (The masses listed here assume 100 percent decomposition of the N2H4 into N2
and H2.) Jeff Maybee of Primex Aerospace has been consulted about the design of a minimalammonia
hydrazine catalyst.
Operation of the system begins by actuating the pyro-valve (or latch valve) at the exit of the gas
JPL Document # 16330
23
tank (Fig. 4.9). Regulated helium gas is then introduced into the struts and torus (controlled by a
series of solenoid valves), with the cooling of the gas over the duration of the fill assumed to be
within acceptable parameters for the cold-rigidized portions of the structure. The initial inflation of
the structure is assumed to have a 5-minute duration (worst case). The next operation will be
pressurization of the solar array booms to sufficient internal pressure (8 psia) to extend and
rigidize, followed by inflation of the antenna envelope to operating pressure. The isolation valve
between the hydrazine tank and gas tank is then opened, allowing the gas tank to serve as a
reservoir for the products of the hydrazine catalyst bed. The gas tank is of sufficient volume to
maintain an approximately one-day supply of make-up gas at nominal conditions, at 5-atm tank
pressure. (See section 6.1 for a description of the full launch sequence).
Further definition of the inflation system is dependent on refined estimates of the operational
requirements. If the expected catalyst bed exit temperature of 800K is within the allowable
temperature range of the spacecraft structural elements, it may, for example, prove feasible to
eliminate the gas tank and use the hydrazine system for both initial inflation and make-up. While
this does not provide a significant mass advantage, the resulting system will have fewer
components. Additionally, the leakage rate at the end of the mission determines the catalyst size
and allowable quiescent period for the science mission, unless the fill operation can be done
concurrently with data acquisition. The current configuration incorporates a catalyst of 0.2 kg,
which is estimated to be sufficient to provide a gas flow rate of 0.6 kg/min.
Potential issues which have yet to be quantitatively addressed include integration of the inflation
and propulsion systems, possible absorption of signals by the inflation gas (especially ammonia),
possible ionization of inflatant gases caused by large electrical potential gradients between the
canopy and reflector (and resultant effect on data collection), and condensation of inflation gases at
minimum-temperature conditions.
Fill
MV
PV
C
A
T SV
F
O
Fill
Fill
SV
PV
MV
MV
F R R
LHZ Tank
Gas Tank
Figure 4.9: ARISE Inflation System: MV = manual valve; SV = solenoid valve; PV = pyro
valve; F = filter; R = regulator; O = orifice
JPL Document # 16330
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4.6 Deployment sequence and Canister design
L’Garde has developed a new flexible enclosure canister as shown in Figure 4.10 and 4.11. The
concept promises significant weight savings over the solid canister designs. The lenticular and
torus are stored inside a membrane container designed to withstand the increased internal pressure
during ascent of the payload. Upon deployment, the top portion of the membrane is released by a
pyrotechnic. The petals open, releasing the lenticular and torus. Some residual gas in the lenticular
is possible (as was experienced in the IAE flight experiment) and the lenticular is expected to
billow out slightly to relieve any internal pressure. After initial deployment the LDDs or L’Garde
Deployment Devices are initiated. The next picture in the sequence shows the LDDs deploying the
struts. Note, the torus and lenticular are still uninflated and suspended between the extending
struts. The next pictures show the struts at full deployment and the torus partially inflated. The
lenticular is still not inflated. The solar array gets deployed followed then by the inflation of the
lenticular structure. Development of the deployment sequence draws extensively on the lessons
learned from the IAE flight experiment.
Release Activation Struts deploy
Figure 4.10. ARISE Canister Concept
Figure 4.11. Inflatable Antenna Deployment Sequence (by TDM Inc.)
JPL Document # 16330
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Figure 4.11. Inflatable Antenna Deployment Sequence (by TDM Inc.) (continued)
4.7 Subreflector description
The basic configuration and RF error budget for the ARISE radio telescope assumes that the
deviations from an ideal surface figure in the primary mirror (or reflector) will be compensated in
large part by changing the shape of the secondary reflector. Present assumptions suggest that this
correction may have to be done as often as once every fifteen minutes, but not at a frequency of
cycles per second. There is work presently underway at Composite Optics Inc. (COI) in tunable
radio frequency reflectors in the one to two meter diameter range. Moreover, work is being done
on SBIR contracts by a number of vendors to solve the actuator requirements of the Next
Generation Space Telescope (NGST). Given this work in progress the assumption is that an
actuated, tunable reflector with the desired optical characteristics may be manufactured at
reasonable cost, by building on the work in progress at COI and NGST.
There are six basic problems to be resolved with the ARISE secondary mirror task. These are:
- Mirror Skin Design
- Actuator Selection
- Strong Back
- Mass
- Software
- Integration & Launch Packaging
Mirror Skin Design
COI has completed a Phase One SBIR for ground based tunable reflectors of comparable size to
ARISE’s subreflector. These reflectors have a composite surface, and are adjusted with screw
jacks on the back surface of the mirror. COI has recently been awarded a phase two SBIR contract
to extend this work. Of particular relevance to the ARISE work is the amount of finite element
modeling which is being done under these contracts, and the confirmation of these FEM models in
full size test mirrors. The models comprise a set of basic design tools which COI can employ, at
modest cost, on a prototype design of the ARISE reflector. A request for contract numbers, and
relevant technical publications is pending with COI, as of this writing.
JPL Document # 16330
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Actuator Selection
To shape the surface of the secondary reflector, a set of cryogenically adapted, precision, linear
motors is required. NGST requires more than 2000 cryogenic actuators, some of which must have
a travel of millimeters, and others of which must have a resolution in the 10’s of nanometers.
NGST is presently funding research at the level of around $800,000 per annum in mechanism
development. These mechanisms are being studied in the Low Temperature Science & Engineering
cryogenic mechanisms laboratory at JPL. From this work there is a high degree of certainty that
actuators suitable for ARISE will be produced. A significant candidate for this actuator is a
cryogenic lead-screw designed by Thermetrex Corporation of San Diego. This actuator is
tentatively scheduled for test at JPL this fall.
Strong Back
The actuators that shape the secondary need a stable, structural reaction surface behind the mirror.
For lightness and strength this should probably not be inflatable, unless it is a self-rigidized
inflatable of some type, perhaps with filled epoxy. A composite strong back structure should work
well. Because COI has experience in understanding the composite mirror loads, they are the logical
company to design a strong back structure that accurately reacts to those loads. The required mass
for such an assembly is not known as of this writing. A small study may be sufficient for COI to
adapt their present SBIR work to the ARISE problem. This should probably be done sooner rather
than later, in case the mass of the strong back is greater than that included in the present ARISE
mass budgets.
Mass
Generating an adequate predictive model on the mass budget for the secondary mirror assembly is
a task that requires some attention at this point. The strong back, actuator, actuator electronics,
cables, supporting structure and the mirror masses are not known at this time to any degree of
specificity.
Software
Various organizations have largely resolved the problem of feeding wave-front error into optical
surface adjustments over a broad range of wavelengths and response frequencies. A brief survey
of the established methods for doing this should be undertaken, in order to select the most
appropriate to be adapted to the needs of ARISE.
Integration and Launch Packaging
Significant economies in mass, and other types of performance advantage can often be achieved by
cleverly resolving the design trades in the integration and packaging task. For instance, it may be
possible to reduce mass, and launch package volume by designing a strong back / mirror surface
system in which a primary and secondary inflatable structure are used. The primary structure
might hold the mirror / actuator assembly in place. An inflatable, epoxy-cure-upon-deployment
strong back could be designed to provide needed strength to the system.
These design trades are often best done at the early phases of system conceptualization. Therefore a
brief integration design-trade study, with a view to mass and volume reduction should be
conducted in the very near future.
JPL Document # 16330
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5. Science Payload
5.1 Science requirements
The ARISE top level science requirements can be summarized as:
1. Source detection: the spacecraft must be able to detect sources that have strength of about
10 mJy at 43 GHz, and about 3 mJy at 22 GHz.
2. Spacecraft - ground telecom data rates: 8 Giga bits per second (Gbps), driven by
sampling at Nyquist rate (2 samples/sec/Hz) and digitizing at 1bit/sample or 2 bits/sample.
3. Observation duration: the spacecraft must be able to observe a single source for 12-24
hours, at one or several frequencies. One coherent integration time is between 15 and 350 seconds.
4. Sampling duty cycle: the fraction of time that science data is gathered during one
observation is at least 70 %. That leaves 30% for other spacecraft duties.
5. Gain variation: the gain in the direction of the source cannot vary more than 2-5% during
one coherent integration time.
Table 5.1 to 5.12 provide a more complete set of default parameters for the ARISE mission,
together with possible ranges for those parameters. Telescope parameters for the Green Bank
Telescope (GBT) have been taken from the GBT web site, while VLBA telescope parameters are
taken from the VLBA web site, with some assumptions made about improved system temperatures
by the time of ARISE launch. The possible ranges of many parameters are educated guesses. It is
unlikely that phase-referencing will be possible at the higher (or any) frequencies for ARISE.
However, the tables include the possibility of achieving equivalently long coherence times by
means of atmospheric calibration at the ground telescopes (using water vapor radiometers, phase
referencing of the ground telescopes, or similar techniques). Finally, in the calculation of spectralline
sensitivities, a channel width of 0.5 km/sec has been assumed in all cases. This was chosen as
a compromise among the various types of spectral-line science that might be done, and can easily
be scaled for other assumptions by the square root of the channel width.
Table 5.1: Radio Telescope
Quantity Nominal Possible Range
Diameter 25 m 15 - 25 m
Structure Inflatable Others
Optics Off-axis Gregorian On-axis; Cassegrain
Sun-Avoidance Angle 30 deg 20 - 50 deg
Pointing Accuracy 3 arcsec 2 - 6 arcsec
Slew Rate 2 deg/min 1 - 4 deg/min
Phase Referencing None 5 - 8 GHz
Surface Accuracy 0.5 mm (target) 0.2 - 1 mm
Corrected Surf. Acc. 0.25 mm 0.2 - 0.5 mm
Table 5.2: Observing System
Quantity Nominal Possible Range
Freq. Coverage 8, 22, 43, 60, 86 GHz No 8 & 60; add 1.6 & 5
Polarization Dual Circular Single Circular
Polarization Purity < 3% 1-6%
Sampling 1 or 2 bit 2 bit
Calibration Accuracy 2% 1% - 3%
IF channelization TBD TBD
JPL Document # 16330
28
Table 5.3: Sensitivity vs. Frequency - 8 GHz
Quantity Nominal Possible Range
Frequency Span 8 - 9 GHz 5 - 9 GHz
Tsys 12 K 8 - 15 K
Aperture Efficiency 0.50 0.4 - 0.6
System Equivalent Flux
Density (SEFD)
130 Jy 75 - 590 Jy
Coher. Time (C=0.9) 350 sec 100 - 2000 sec
Data Rate 4 Gbit/sec 1 - 4 Gbit/sec
Table 5.4: Sensitivity vs. Frequency - 22 GHz
Quantity Nominal Possible Range
Frequency Span 21 - 23 GHz 18 - 23 GHz
Tsys 16 K 12 - 25 K
Aperture Efficiency 0.38 0.3 - 0.5
SEFD 240 Jy 130 - 1280 Jy
Coher. Time (C=0.9) 150 sec 60 - 1000 sec
Data Rate 8 Gbit/sec 1 - 8 Gbit/sec
Table 5.5: Sensitivity vs. Frequency - 43 GHz
Quantity Nominal Possible Range
Frequency Span 42 - 44 GHz 40 - 45 GHz
Tsys 24 K 20 - 35 K
Aperture Efficiency 0.24 0.2 - 0.35
SEFD 560 Jy 320 - 2700 Jy
Coher. Time (C=0.9) 60 sec 20 - 400 sec
Data Rate 8 Gbit/sec 1 - 8 Gbit/sec
Table 5.6: Sensitivity vs. Frequency - 60 GHz (single dish only)
Quantity Nominal Possible Range
Parameters TBD TBD
Table 5.7: Sensitivity vs. Frequency - 86 GHz
Quantity Nominal Possible Range
Frequency Span 84 - 88 GHz 80 - 90 GHz ?
Tsys 45 K 30 - 80 K
Aperture Efficiency 0.08 0.08 - 0.2
SEFD 3200 Jy 850 - 16,000 Jy
Coher. Time (C=0.9) 15 sec 5 - 100 sec
Data Rate 8 Gbit/sec 1 - 8 Gbit/sec
JPL Document # 16330
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Table 5.8: 7-sigma continuum sensitivity to 1 VLBA antenna
Quantity Nominal Possible Range
8 GHz 1.9 mJy 0.6 - 15 mJy
22 GHz 4.5 mJy 1.3 - 46 mJy
43 GHz 15 mJy 4.4 - 162 mJy
86 GHz 120 mJy 25 - 1400 mJy
Table 5.9: 7-sigma continuum sensitivity to GBT
Quantity Nominal Possible Range
8 GHz 0.4 mJy 0.1 - 3.6 mJy
22 GHz 0.8 mJy 0.2 - 8.3 mJy
43 GHz 2.5 mJy 0.7 - 28 mJy
86 GHz 26 mJy 5.5 - 300 mJy
Table 5.10: 7-sigma spectral line sensitivity to VLBA antenna (0.5 km/s channel)
Quantity Nominal Possible Range
8 GHz 0.5 Jy/ch 0.2 - 2.2 Jy/ch
22 GHz 1.0 Jy/ch 0.3 - 3.8 Jy/ch
43 GHz 2.5 Jy/ch 0.7 - 9.5 Jy/ch
86 GHz 14 Jy/ch 2.9 - 57 Jy/ch
Table 5.11: 7-sigma spectral line sensitivity to GBT (0.5 km/s channel)
Quantity Nominal Possible Range
8 GHz 0.1 Jy/ch 0.04 - 0.5 Jy/ch
22 GHz 0.2 Jy/ch 0.06 - 0.7 Jy/ch
43 GHz 0.4 Jy/ch 0.1 - 1.6 Jy/ch
86 GHz 3.1 Jy/ch 0.6 - 12 Jy/ch
Table 5.12: Additional mission information
Quantity Nominal Possible Range
Launch Date 2008 2007 - 2012
Lifetime 3 yr 2 - 5 yr
Comm. Link 38 GHz 80 GHz or optical
Tracking Stations 5 4 - 7
5.2 Receivers/Amplifiers
The ARISE receiver design is critical to the final performance of the instrument. By providing the
lowest noise possible, requirements on the size and performance of the primary mirror can be
bound to achievable goals. Cryogenic InP High Electron Mobility Transistors (HEMTs) provide
the lowest possible noise for receivers from 1-100 GHz operating at temperatures above 4 K. In
JPL Document # 16330
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addition InP HEMT transistors operate with the lowest power dissipation of any three terminal
device, which is critical for thermal load on the cryocooler.
The baseline ARISE receiver has channels at 8, 22, 43 and 86 GHz. The receiver front end is
cooled to 20 K. The front end (Figure 5.1) is comprised of an antenna, orthomode transducer and
an InP HEMT amplifier. The front end is nominally designed to have a noise figure of 5 times
quantum limited noise at all four frequencies. This noise temperature has already been achieved at
8, 22 and 43 GHz using InP HEMT amplifiers with discrete transistors. This goal at 86 GHz is
expected to be met by a cryogenic amplifier program at JPL and is consistent with the goals of the
ESA’s Planck Surveyor, Low Frequency Instrument. The use of InP monolithic millimeter-wave
integrated circuit (MMIC) technology allows for state-of-the-art performance and ease of
integration at 86 GHz. This will be crucial should the adaptive array be utilized on ARISE.
Parameter 5 GHz 22 GHz 44 GHz 86 GHz
Noise (K) 8 12 19 39
Bandwidth
(GHz)
2244
Cryo power
(mw)
64 23 15 8
The cryogenic portion of the radiometer front end is connected to a warm back end via stainless
steel waveguide or coaxial cable. The signals are then passed through an image reject filter and
mixed down to an IF bandwidth of DC-4 GHz. The mixer will utilize a phase locked local
oscillator (PLLO) with phase locking derived from a stable crystal oscillator. Should an adaptive
array be implemented, each feed in the array will have its own front end and mixer, the IF signal
will be passed into an array processor which will trim the phase of the PLLO on each element
individually and adjust the IF amplifier gain. The output of the array processor will be a single IF
channel with a “clean” effective beam.
Figure 5.1: ARISE Receiver Schematic.
Table 5.13 Performance of the ARISE Receivers.
8 43
JPL Document # 16330
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Following the array processor, the IF signals are sorted by frequency and polarization and
digitized. A tone provided by a ground based source is digitized along with the signal to provide an
accurate phase reference for the signal.
Two approaches towards digitization may be taken. The first option is to build on the current
VLBA digitizers which multiplex many 32 MHz A/D converters to build a larger bandwidth. The
advantage of this scheme is that it takes advantage of the current VLBA equipment, potentially
reducing overall program costs. A second approach is to utilize modern high speed A/D converters
operating at frequencies in excess of 1 GHz. The advantage to this technology is a greatly reduced
digitizer mass and power, but the development costs to populate VLBI telescopes could be
significant.
5.3 RF adaptive compensation
At all the operating frequencies, the radiation performance of ARISE has been evaluated using a
vector diffraction computer program, which employs Physical Optics on both the main reflector
and the subreflector. For the operating frequencies at 43 GHz and 86 GHz, array feeds are utilized
to electronically compensate for the performance deterioration caused by surface distortions and
beam pointing error. At both operating frequencies, a 19 element array feed is used with 0.86 λ
inter-element spacing. Investigations into adaptive methods to optimally combine the signals
received by the individual array elements are currently under way. Utilization of circularily
polarized compensation is also under investigation, and it is expected that although results will not
change, hardware and software implementation will be more complex.
5.3.1 Feed layout and array configuration
1.4cm
1.4cm
3 1 2
4
5
Figure 5.2: Layout of the single feed and array of horn feeds for the five different operating
frequencies. The numbers refer to 1 = 86 GHz, 2 = 43 GHz, 3 = 22 GHz, 4 = 8 GHz and 5 = 4.85
GHz (Size of scaled square corresponds to size of the 86 GHz feed).
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The layout of the horn feeds and array of horn feeds in the focal plane of the reflector are displayed
in Figure 5.2. The numbers in this figure refer to the operating frequencies of the respective feed,
i.e.
1 = 86 GHz
2 = 43 GHz
3 = 22 GHz
4 = 8.0 GHz
At both 43 GHz and 86 GHz a 19-element array feed is used. Both array feeds are hexagonal
based. The feed geometry is displayed in Figure 5.3. The usage of a 37-element array feed is also
contemplated. This array feed is similarly hexagonal based and its feed geometry is displayed in
Figure 5.4. For the RF performance shown in the following paragraphs, an array inter-element
spacing of 0.86 λ is assumed. This array spacing is still a subject of ongoing research.
Array Geometry 19 Elements
<Element #>
y = 0
x = 0
1
2
3
4
5
6
7
8
9
10
11
12
13
14
15
16
17
18
19
Figure 5.3: Geometry of the hexagonal based 19-element array feed. The array spacing is at 0.86
λ, which is 6 mm at 43 GHz and 3 mm at 86 GHz.
5.3.2 Surface distortions and un-wanted beam tilt
Due to material, manufacturing and deployment imperfections, limitations and errors, surface
distortions are introduced in the primary reflector. Because of the unique characteristics of the
inflatable membrane structure, these distortions are slowly varying in nature. In order to accurately
simulate the RF system performance, the assessment of these distortions in terms of their effects
on the radiation characteristics is imperative.
JPL Document # 16330
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Array Geometry 37 Elements
<Element #>
y = 0
x = 0
1
2
3
4
5
6
7
8
9
10
11
12
13
14
15
16
17
18
19
20
21
22
23
24
25
26
27
28
29
30
31
32
33
34
35
36
37
Figure 5.4: Geometry of the hexagonal based 37-element array feed. The array spacing is at 0.86
λ, which is 6 mm at 43 GHz and 3 mm at 86 GHz.
As an analytic distortion model, a functional dependency of the form:
z = τ * ρ3
* sin(3 * ϕ)
was considered. (ρ, ϕ) are the coordinates in a polar coordinate system, τ denotes the center-topeak
height of the distortion. A surface plot of this analytic distortion model is displayed in Figure
5.5. A surface RMS value of 0.5 and 1 mm was considered, which resulted in a center-to-peak
height of 1.5 and 3 mm in the distortion model described above.
A surface plot of the most recently supplied discrete surface distortion data by L'Garde is displayed
in Figure 5.6. This distortion has a surface RMS value of approximately 1 mm.
A further source of RF performance degradation is un-wanted beam pointing. A beam pointing
error of only one beamwidth reduces the directivity and antenna efficiency significantly. At higher
frequencies the beamwidth becomes more narrow, which makes the accurate pointing more
difficult and necessitates the usage of an array feed and corrective feed array excitation to achieve
the required RF performance.
JPL Document # 16330
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Figure 5.5: Surface plot of the analytic surface distortion model. The center-to-peak height is 3
mm, which yields a rms of approximately 1 mm.
Figure 5.6: Surface plot of the discrete surface distortion data as supplied by L'Garde on 04/06/98.
5.3.3 RF performance using single feeds
At first the RF system performance is evaluated using single feeds at each of the five operating
frequencies. This approach is less costly and easier in its implementation complexity. The key
ARISE RF performance parameters for an undistorted and distorted surface (analytical distortion
model) using single feeds at each frequency are displayed in the following table. For the analytic
distortion model, an RMS value of 1 mm has been considered. Additionally at the operating
frequencies 43 GHz and 86 GHz, a surface RMS value of 0.5 mm has been investigated.
JPL Document # 16330
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Frequency [GHz] Configuration D [dB] AE [%] BW [deg.]
8.0 ideal 65.52 81.0 0.110
distorted (1 mm) 65.19 75.3 0.110
22.0 ideal 74.30 81.0 0.036
distorted (1 mm) 71.95 47.3 0.045
43.0 ideal
distorted (0.5 mm)
80.19
77.87
81.0
48.4
0.018
0.022
distorted (1 mm) 73.79 18.2 0.033
86.0 ideal
distorted (0.5 mm)
86.14
79.82
81.0
19.0
0.009
0.017
distorted (1 mm) 76.66 9.1 0.019
Table 5.14: RF performance of ARISE, using a single feed at each of the operating frequencies.
In Table 5.14, D denotes the directivity in dB, AE the antenna efficiency (including taper and
illumination efficiency) and BW the half-power beamwidth in degrees. Note that the efficiency
referred to here includes only taper and illumination efficiency. Other efficiency numbers due to
losses caused by effects such as mismatch, feed network, finite surface conductivity, local surface
rms, blockage or polarization mismatch are not considered in the antenna efficiency. The ultimate
efficiency will be somewhat lower than the value recorded in this table.
It is apparent from the previous table, that the deteriorating effects of the surface distortions
increase with frequency. While the losses in directivity and, hence, antenna efficiency are
acceptable up to a frequency of 22 GHz, the losses at 43 GHz and 86 GHz are too severe for the
RF system performance requirements assuming a RMS value of 1 mm. New ways to achieve the
requirements are needed.
5.3.4 RF performance using array feeds
At both 43 GHz and 86 GHz, array feeds are used to compensate for surface distortions and beam
pointing errors. In the following, the RF performance at 43 GHz and 86 GHz is discussed using
an array feed as displayed in part 5.3.1 of this section. At first, analytic surface distortions as
displayed in Figure 5.5 are considered. In the following table, the three cases 'ideal' (no surface
distortions), 'distorted' (analytic surface distortion model) and 'compensated' (using corrective
feed array excitation) are considered.
For the discrete surface distortion data as supplied by L'Garde and shown in Figure 5.6, the RF
performance at 43 GHz is investigated. With these surface distortions, the directivity reduces to
72.01 dB, where the beam tilts by 0.029 degrees. With a 19-element array as displayed in Figure
5.3 for distortion compensation, the directivity changes to 70.27 dB and the beam tilt reduces to
0.025 degrees. Using a 37-element array as displayed in Figure 5.4, the directivity increases to
73.31 dB and the unwanted beam tilt is fully compensated for. To avoid the usage of a 37-element
array, subreflector shaping in combination with a 19-element array compensation is currently under
investigation.
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Frequency [GHz] Config. D [dB] AE [%] BW [deg.]
43.0 (RMS=0.5mm) ideal 80.19 82.40 0.018
distorted 78.11 51.07 0.022
compensated 78.38 54.32 0.022
43.0 (RMS=1mm) ideal 80.19 82.40 0.018
distorted 74.34 21.42 0.032
compensated 75.36 27.12 0.030
86.0 (RMS=0.5mm) ideal 86.18 81.87 0.009
distorted 80.34 21.32 0.016
compensated 81.67 29.01 0.015
86.0 (RMS=1mm) ideal 86.18 81.87 0.009
distorted 77.17 10.29 0.019
compensated 78.63 14.41 0.018
Table 5.15: RF performance of ARISE at 43 GHz and 86 GHz, using 19-element array feeds with
corrective feed array excitation coefficients.
For the beam pointing error, an unwanted beam tilt in the amount of the half-power beamwidth is
considered. At an operating frequency of 43 GHz, the half-power beamwidth is 0.018 degrees.
Given a beam tilt of the same amount (0.018 deg.), the directivity and antenna efficiency at
boresight without array compensation reduces to 67.13 dB and 4.08%, respectively. With the
array compensation, the beam tilt is reduced to 0.006 degrees and the maximum directivity at that
location is at 79.1 dB. For an unwanted beam tilt of 0.012 degrees, the array compensation can
fully compensate for the tilt. The directivity for the uncompensated and the compensated case are
75.12 dB and 79.64 dB, respectively.
5.3.5 Representative farfield patterns
The operating frequency for the following four cases is at 43 GHz. In Figure 5.7 a), the beam
contour pattern of an undistorted ARISE reflector surface is displayed. In Figure 5.7 b), the
distorted beam contour pattern is displayed using the discrete surface distortion data supplied by
L'Garde (see Figure 5.6). A single feed is used in these two cases.
In Figure 5.8 a), a 19-element array is used to compensate for the surface distortion effects. In
Figure 5.8 b), a 37-element array is used to compensate for the surface distortion. The
improvement using array feed compensation with 19 elements and especially 37 elements is clearly
visible in these figures.
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a) b)
Figure 5.7: Beam contour pattern of ARISE at 43 GHz. a) No surface distortions. Single feed. b)
Discrete surface distortion model as shown in Figure 5.6. Single feed.
a) b)
Figure 5.8: Beam contour pattern of ARISE at 43 GHz. Discrete surface distortion model as
shown in Figure 5.6 a) Feed array compensation using a 19-element array feed. b) Feed array
compensation using a 37-element array feed.
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5.4 RF system performances
One of the most critical parameter to assess on ARISE is the overall antenna efficiency. For a 25-m
diameter aperture, the science requirements ask for a 7 σ sensitivity of about 3 mJy at 22 GHz and
about 10 mJy at 43 GHz. It is then desirable that the antenna efficiency be at least 0.55 at 22 GHz,
at least 0.2 at 43 GHz, and the highest possible at 86 GHz. The RF adaptive compensation scheme
described above allow for an increase in antenna efficiency that includes distortions of the main
reflector, aperture taper, feed spillover and polarization efficiencies. However, other losses should
be taken into account. A preliminary assessment of these other system efficiencies is summarized
in Table 5.16 for each frequency of interest. The canopy transmittance is based on testing of a 0.5-
mil CP-1 sheet coated with 100 Ang. of ITO, which is currently the candidate material. The
reflector material is a 0.5-mil sheet of Kapton coated with Aluminum. The pointing error is an
estimate from the spacecraft and antenna ACS analysis.
Efficiency 8 GHz 22 GHz 43 GHz 86 GHz
Adaptive compensation (computed)
Antenna distortions
Aperture taper
Feed spillover
Polarization
0.753
(Single feed)
0.473
(Single feed)
0.271 0.144
RF path attenuation
Shadowing (struts + S/C)
Canopy transmittance (twice)
Meteoroid shield (twice)
Reflector reflectance
Surface local rms (specular)
Feed displacement
0.94
0.932
0.952
0.98
0.98*
0.98*
0.94
0.912
0.952
0.98
0.98*
0.98*
0.94
0.882
0.952
0.98
0.98*
0.98*
0.94
0.852
0.952
0.98
0.98*
0.98*
Pointing error 0.98* 0.98* 0.98* 0.98*
Surface ohmic efficiency 0.99* 0.99* 0.99* 0.99*
Feed network loss 0.95* 0.95* 0.95* 0.95*
Margin 0.97 0.97 0.97 0.97
Total efficiency 0.46 0.28 0.15 0.07
Table 5.16: Projection of the overall aperture efficiency *: estimated efficiency
The prediction of the antenna distortions results from an ACS/dynamics, structural analysis
(described in the inflatable antenna section and spacecraft design section) and thermal steady state
worst case scenario analysis. The amplitude of the distortions are summarized in Table 5.17. More
work needs to be done to assess/confirm these antenna distortions.
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Distortions type Amplitude (peak
to valley) (mm)
RMS
(mm)
Comments
Manufacturing /
Wrinkles
3 1 Predicted
Dynamics < 0.1 Reaction Wheels effects
Thermal TBD Not compensated - Steady
ESD Lofting TBD Unknown at this stage
Make up gas waves TBD Unknown at this stage
Total
Table 5.17: Projection of the inflatable antenna distortions
Assuming the aperture efficiencies stated in Table 5.16, the science system performances then
provide a 7 σ sensitivity summarized in Table 5.18.
Frequency 8 GHz 22 GHz 43 GHz 86 GHz
ARISE diameter
ARISE efficiency
ARISE Tsys
25 m
0.46
12 K
25 m
0.28
16 K
25 m
0.15
24 K
25 m
0.07
45 K
VLBA diameter
VLBA efficiency
VLBA Tsys
25 m
0.72
30 K
25 m
0.52
60 K
25 m
0.36
80 K
25 m
0.15
100 K
Data rate 4 Gbps 8 Gbps 8 Gbps 8 Gbps
Coherence time 350 s 150 s 60 s 15 s
7 σ limit 2.0 mJy 5.2 mJy 19.0 mJy 130 mJy
Table 5.18: ARISE Science System Performance Projection
Sensitivities of the antenna efficiency to the final detection threshold is under investigation.
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6. Spacecraft Description
6.1 Spacecraft configuration
Star Tracker
Science Instruments
8, 22, 43, 60, 86 GHz receivers
GPS receiver (1 of 2)
Subreflector, 1.6 m diameter
Radiators (1 of 2)
1.8 x 0.55 m
2 m
1.3 m
0.4 m 1.8 m
Main Thruster,
450 N LEROS-1C
Solar Arrays
(a la ITSAT)
2.0 m x 6.7 m
ACS thrusters (20 N)
ACS thrusters (0.9 N)
Telecom antenna, 1.2 m diameter
(KA-band)
Inflatables Canister
STOWED CONFIGURATION
Figure 6.1: ARISE spacecraft in the stowed launch configuration
The spacecraft design is based on an octagonal shaped bus 1.3-m large and 2-m high. Figure 6.1
shows a conceptual external configuration of the ARISE spacecraft. Figure 6.2 shows the
conceptual layout of the interior of the spacecraft bus. The octagonal structure supports the
inflatable antenna canister on the top, as well as the inflatable solar arrays on two of the side
panels. The other panels support a deployable 1.2-m diameter RF Ka-band telecom antenna, a
deployable 1.6-m diameter secondary (sub-) reflector, and various other spacecraft equipment
(ACS thrusters, GPS receivers, star scanners, radiators, omni-antennas...). The science receivers
described in the science subsystem section are located at the focal plane and on the same panel as
the sub-reflector and below the canister in a way that no blocking of the receivers occurs. The
canister was designed to be part of the bus structure as much as possible (to reduce its structural
mass) and to minimize blocking/shadowing of the main reflector.
The spacecraft volume and maximum dimensions were mostly driven by the Delta II 7925 9.5-ft
diameter three-stage configuration fairing. The interior dimensions of the fairing are 2.5-m
diameter at the base (same diameter for a height of 2-m) and about 4.6-m high. The canister
diameter was constrained by the width of the shroud. In the current design, there is about a 0.3-0.4
m radial margin with respect to the shroud for the lower part of the spacecraft (everything below
the canister).
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PMAD
Li-ion Batteries
Solar Array Drive
Sterling coolers C&DH
Telecom X-band
transponder
Telecom Ka-band
NTO Catalyst bed
Reaction wheels
(Teldrix DR50)
Inflation electronics
NTO tank
Hydrazine tank
Main Engine
(LEROS 1-c)
Electronics deck
Science instruments
electronics Main bus structure
Cryo electronics
Figure 6.2: Inside layout of the ARISE spacecraft
The interior layout of the spacecraft was driven by the Delta II 7925 Cg requirement, which must
be located about 1.2 m above the separation plane (see Fig. 6.2). The fairing separation plane
corresponds to the bottom plane of the spacecraft bus where the adapter will be located. The lower
third of the spacecraft bus contains mainly the propulsion module with the Hydrazine, Nitrogen
Tetroxide, Xenon and pressurant tanks, feed systems, main engine and mounting, inflation catalyst
bed and various hardware associated with the inflation system. About 295 kg of fluids will be
initially loaded, and about 128 kg of propulsion dry mass (no contingency) will be mounted in this
first third of the spacecraft bus. The middle third will house the electronics deck, with the Telecom,
Data system, Power, Attitude Control and Science hardware and electronics. This deck also
includes the two cryocooler stages, with the Sorption cooler mounted at the bottom of the inflatable
antenna canister. No or little effort was done to integrate the electronics with the structure as per the
Lookheed Martin Multifunctional Structures bus design. By 2008, it is to be anticipated that such
an integrated bus will be current technology and thus the spacecraft design will have to be revisited
to take this technology into account (a projected 20% reduction in total spacecraft mass could be
applicable then). The top third of the spacecraft bus holds the inflatable antenna canister which is
integrated with the bus structure.
About 6.9 hours after launch, a perigee raise maneuver will occur at the GTO (Geo Transfer Orbit)
apogee. Then several sequences of deployment will happen. First, due to its large size, the
inflatable antenna will be deployed. This deployment will be controlled and will take between 5 and
20 minutes. The struts will be inflated first, then the torus. The struts and torus will be rigidized
through a thermal phase change of the material (cold rigidization). To simplify the inflation system,
the inflatable solar arrays will be deployed (on 2 wings) next. This scenario enables the
combination of the antenna inflation system, solar arrays inflation system and the attitude
control/propulsion system, thus reducing system dry masses. Only once the solar arrays are in
place does the reflector/canopy assembly get inflated. The spacecraft will be running on batteries
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until solar arrays deployment and will be telecommunicating with the Earth with two
omnidirectional antennas. The third deployment will be the sub-reflector one. A rigid astro-mast
type arm will be used to carry the sub-reflector to about 3.6 m from the spacecraft. A gimbal
system at the end of the mast will then align the sub-reflector with the main reflector. The last
deployment will be that of the 1.2-m diameter telecom antenna. This antenna needs to be deployed
downward with respect to the spacecraft bus in order to get a clear half-space field of view and also
to minimize coupling of Ka-band telecom antenna with the main reflector. A gimbaling system will
allow the antenna to rotate and cover a whole half space.
A challenge in the spacecraft design was to take into account all the pointing and field of view
requirements. During science observations, the main reflector can be pointed anywhere in the sky
except for a 30 deg. cone around the Sun. At the same time the solar arrays must be pointed toward
the Sun to provide the 2.4-kW needed (see power budget in section 6.2), and the Ka-band telecom
antenna must be pointed toward the Earth for continuous science data downlink. To achieve these
requirements, the solar arrays will be one-axis gimbaled and the telecom antenna is deployed and 2
DOF gimbaled to cover a half-space. At this time it is believed that this configuration should allow
for almost continuous coverage of the telecom ground stations but a more thorough analysis should
be done to evaluate the effective coverage.
6.2 System description, mass and power budgets
A top level mass budget of the ARISE spacecraft is summarized in Table 6.1. A more detailed
mass budget can be found in Appendix A. A 30% mass contingency was applied to the spacecraft
dry mass. The propulsion module was sized using the launch vehicle injected mass. The main
GTO perigee raise maneuver is done with the NTO/Hydrazine 450 N Leros 1-C, which features an
Isp of 325 seconds. A ∆V of about 380 m/s is achieved. Height (8) 22 N thrusters will be used for
main burn trajectory correction, and height (8) 0.9 N thrusters will be used for coarse attitude
control maneuvers.
Subsystem Mass (kg)
Inflatable antenna 192
Telecom 32
C&DH 13
Power 142
ACS 82
Thermal Control 159
Structures/Mechanisms 204
Propulsion (+ACS) system 128
Science instruments 81
Spacecraft dry mass 1033
Contingency (30%) 310
Propellants/Fluids 295
Launch Vehicle adapter 46
Total spacecraft mass 1684
Launch vehicle capability (@ i=36 deg.) 1691
Table 6.1: ARISE spacecraft mass budget
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Twelve (12) mini ion thrusters (3-cm diameter, see description in Appendix D) will be used on the
inflatable antenna torus to overcome and correct for solar pressure torques. These mini-ion
thrusters still need to be developed, and 3 sets of the more mature Field Emission Electric
Propulsion (FEEPs) could be used instead for about the same dry mass. One concern though with
the FEEPs is that they use liquid metal as propellant and therefore contamination of the canopy and
main reflector might be an issue. However, more analysis needs to be done to determine the
implications of having ion thrusters on the inflatable antenna torus, both on an ACS and structures
points of view. The attitude determination of the spacecraft will be done with star trackers, sun
sensors and 2 GPS receivers. Fine pointing will be done with reaction wheels. It is to be
anticipated that a closed loop between the science receivers (feed array) and the ACS system will
have to be designed in order to achieve the 3-50 arcsec pointing requirements of the main reflector.
The science data requirements are quite demanding (8 Gbps) and an early trade-off between optical
and RF downlinks was done. In view of the technology development programs, it was decided
that the RF system was a “safer” candidate and therefore chosen as the telecommunication system
for the ARISE spacecraft. This choice could be revisited later on. Thus, a 1.2-m diameter Ka-band
antenna transmitting 30 W RF will be used to downlink science data at a rate of 8 Gbps. Both
polarizations and a high order modulation technique such as the Quadrature Amplitude Modulation
will enable the data to be transmitted within the tight 1 GHz bandwidth available at Ka-band. A
parallel X-band transponder will be used to uplink commands and valuable time information for
VLBI processing. X-band receivers are two patch antennas that will cover near 4π steradian
coverage.
Mode /
Subsystem
Launch
+ postlaunch
Orbit
insertion
All
deployments
Science Stand
by,
slewing
Eclipses
ACS 133.5 133.5 133.5 228.5 228.5 133.5
Propulsion 60 60 70 520 10 10
C&DH 9 9 9 12 9 9
Inflatable antenna 82 82 82 5 5 5
Power 20 20 20 72 40 40
Mechanisms 0 0 80 80 80 50
Telecom 36 36 36 303 36 36
Thermal control 70 70 70 450 70 70
Science 0 0 0 75 0 0
Subtotal (W) 410.5 410.5 500.5 1745.5 478.5 353.5
Contingency
(30%)
123.1 123.1 150.1 523.7 143.5 106.1
Battery recharge 125 125
Total (W) 533.6 533.6 650.6 2394.2 747 459.5
Source Primary
battery
Secondary
battery
Primary
battery
Solar
array
Solar
array
Secondary
battery
Duration (hrs) 6.8 0.5 2 0.75
Table 6.2: Power budget in Watts per mission modes
The power sources encompass a set of primary Li/SO2 batteries for post-launch activities (about 5
kWhr), a set of secondary Li-ion batteries (about 350 Whr, with 125 W of recharge power) for
power generation during eclipses and a 2-wing inflatable solar array for main power generation.
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The array is sized to provide 2.4 kW during science observation (most constraining mode). The
inflation system of the solar array is integrated with the main reflector inflation system and
propulsion/ACS hardware to reduce dry mass. Power demand per mission modes is summarized
in Table 6.2 (see Appendix B for more details). A 30% power contingency was used on all modes.
The larger power consumers are the cryo-coolers, the reaction wheels, the telecom 8 Gbps
downlink and the mini-ion engines. Also, since the power requirements are large during the
science mode, no science will be done during eclipses. Details on all subsystems are given next
sections.
6.3 Gain and observation duration budget
As part of the science requirements, it was prescribed that for valuable science to occur, the gain of
the antenna should not vary more than 2-5% during the sampling period and that the science data
acquisition should take at least 70% of the observation duration. In an attempt to address both
requirements, lists of potential perturbation during sampling and spacecraft duties during
observations were established. Potential perturbations include ACS reaction wheels, cryocooler,
thermal and dynamics, lenticular pressure maintenance, and ESD Lofting. Spacecraft duties are
summarized in Table 6.3. Both list are evolving as our understanding of the spacecraft, mission
and antenna improves.
Task Frequency (per orbit) Total Duration (min)
Earth occultation 0-1 30-40
Antenna calibration
Bright source pointing
Antenna stabilization
Subreflector focusing
Adaptive array calibration
1-2
1-2
1-2
1-2
2-4
TBD
TBD
TBD
Source data acquisition
Target source pointing
Antenna stabilization
1-2
1-2
2-4
TBD
New source pointing 1 19
Telecom ground station switching 1-3 5-15
Telecom link establishment 1-3 2-6
Reaction wheels unloading 32 32
Cryos set-up 1-2 TBD
Antenna gas maintenance 3-? TBD
Miscellaneous
Effective total
Total required 30% of orbit
TBD (120+)
245
RF data acquisition 70% of orbit 571
Table 6.3: Observation duration timeline
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6.4 Spacecraft Data Flow
6.4.1 Avionics
Requirements and Assumptions
The Command & Data Subsystem (CDS) for the ARISE spacecraft is required to operate for a
primary mission life of 3 years. The dual string block redundant design will transfer real time
science data to the Telecom subsystem at a rate of 8 Gbps. The CDS will receive periodic uplink
commands at rate 2 kbps. An 8 GHz uplink tone will be distributed to the science instrument. The
electronics must operate through an equivalent radiation environment of 315 krads behind 100 mils
of aluminum. The mass storage element in the CDS is not required to store or process the high
speed science data.
Design Implementation and New Technology
The CDS required functions are performed by either CDS block redundant strings. Redundancy
provides a high level of reliability to meet the primary and extended mission high speed data
throughput performance requirements. All of the key elements in the CDS design are new
technology.
Real time science data is transferred to the Telecom subsystem via four channels simultaneously.
Each FireWire IEEE 1394.B channel will operate at 2 Gbps. High speed First In First Out
(FIFOs) registers insert header data and State Of Health (SOH) data into the downlink data frames.
The FIFOs may be provided by UTMC or Honeywell. Uplink commands are processed at 2 kbps.
The Lockheed Martin PowerPC 750 processor verifies and processes uplink commands, stores
time tag commands & GPS data, controls the spacecraft fault protection, routes the science data,
distributes the spacecraft time and collects SOH data. Lockheed Martin PowerPC 405
microcontrollers may be used as needed. The flight code and SOH data are stored in flash nonvolatile
memory. Redundant low power serial busses provide an interface to control and monitor
the health of the other subsystems. The flight software code is written in ANSI C or C++. The
ACS pointing and control algorithms are supported in the CDS software.
The 14 Multi-Chip Modules (MCMs) in the CDS design have a mass of approximately 4 kg. Each
CDS string dissipates 9 watts. One CDS string is active at any given time, while the other string is
in a cold sparing mode. Commercial rad-tolerant electronics are shielded behind 100 mils of
Tantalum. The shielding mass is 8.7 kg. The effective radiation environment is 19.5 krads Total
Ionizing Dose (TID). The electronic devices are immune to Single Event Latch-up (SEL) and
immune to Single Event Upset (SEU) to 75 MeV/mg-cm2
.
CDS Avionics Caveats
The IEEE 1394.A FireWire serial bus is presently being developed by the X2000 Team.
Their plan is to provide a fixed rate (100 Mbps) data bus design. Although this does not meet the
needs for the ARISE mission, the X2000-2 Team could leverage from the X2000 design team
effort to develop an IEEE 1394.B FireWire serial bus. The advanced FireWire design would
operate up to 3.2 Gbps. Both FireWire designs require some modification to provide electrical
isolation between functional subsystems. Commercial high speed rad-tolerant electronics will be a
challenge to develop.
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6.4.2 Telecommunications
Requirements
The ARISE spacecraft will orbit around the Earth in a LEO orbit with perigee of 5000 Km and
apogee of 40,000 km. The telecom system needs to downlink data at up to 8 Gbps with bit error
rate (BER) less than 10-4. The observed data is not stored on the spacecraft and needs to be
downlinked immediately. The spacecraft is expected to perform science measurements
approximately 70% of the time. Link availability is required to be greater than 70 % during
observation. Dedicated ground receiving stations are needed. The telecom system also needs to
support a command link at 2 Kbps or less and a housekeeping telemetry link at 2 Kbps or less.
Two-way Doppler measurements need be performed for accurate time-stamping of the data.
Candidate Systems
The candidate telecom system designs consist of an X-band transponder for command,
housekeeping, and two-way Doppler tracking and a separate high data rate downlink system. Both
optical and RF systems have been considered for the high rate downlink.
An optical system was initially recommended because optical systems do not have to meet any
spectral usage requirements. The candidate system employs four wavelength-multiplexed lasers,
each with output power of 2 W. This is designed to match the four channels of radio science data.
The aggregate output is fed to a 30 cm telescope on the spacecraft. With the help of a laser beacon
from the ground receiving site, the spacecraft telescope transmit to the ground station which has a
one-meter telescope. Initial calculation shows that the lasers can provide over 7 dB of link margin.
The high rate RF system needs to operate in the 37-38 GHz spectrum allocated by the FCC for
space VLBI missions. To meet the 8 Gbps downlink requirement with 1 GHz of bandwidth, the
candidate RF system implements two 4 Gbps links using left-handed circular polarization (LHCP)
and right-handed circular (RHCP) polarization. The high rate RF system transmits at 30 W RF in
each polarization to a 34-m DSN antenna. The link margin is about 6 dB.
Optical system offers excellent capability that will greatly benefit ARISE. The ARISE preproject
will work closely with the optical communications technologists to monitor the progress of the
technology development. It is hoped that the optical telecom technology will become sufficiently
matured and that there will be enough operating experience for ARISE to revisit optical
communication systems at a later date.
Current baseline
The current telecom design consists of an X-band transponder for command, housekeeping, and
two-way Doppler tracking and a Ka-band system at 37 GHz band for high data rate downlink.
The X-band transponder design is based on the Spacecraft Transponding Modem (STM). Two Xband
patch antennas are needed for near 4-π steradian coverage. A 0.5 W SSPA provides sufficient
downlink margin. An ovenized oscillator is included to provide accurate Doppler measurement.
The ROM cost is $ 3.3 M. The cost, however, does not include DSN support. DSN cost is
included in a separate ground systems development and operation cost estimate.
With 1 GHz of bandwidth at 37-38 GHz, only high order modulation techniques such as
quadrature amplitude modulation (QAM) can be considered to support the high downlink data rate.
A block diagram of the high data rate transmitter is shown in Figure 6.3. 256-QAM is chosen for
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the current baseline. Each transmitted symbol is selected from one of 256 possible waveforms and
represents eight bits of information. As alternates to the current baseline of 256-QAM, it is
possible to use square-root raised cosine filtering to increase the number of symbols per hertz, thus
allowing the use of smaller QAM constellation such as 16-QAM and 32-QAM which are less risky.
It is also worthwhile to investigate other spectrally efficient modulations such as GMSK and
FQPSK which allow the use of power-efficient non-linear ampliers on the ARISE spacecraft, but
require more spectrum. All of the above options will be examined more closely in future work.
Thirty (30) watts of RF power is needed through a gimbaled 1.2 m high-gain antenna to support
up to 4 Gbps using 34-m DSN stations with 6 dB of link margin. A link budget is shown in Table
6.4. A shaping filter at the transmitter is needed to meet FCC’s spectral usage requirements. This
filter, however, introduces intersymbol interference (ISI) which corrupts the transmitted signal.
An equalizer is needed at the ground receiver. The maximum data rate depends on the FCC
requirements and the complexity of the shaping filter and equalizer. No channel coding is used.
Since the spacecraft also carries a GPS receiver, the gimbaled antenna is expected to be able to
point at the receiving DSN sites without the aid of a beacon from the ground.
The throughput of the 1 GHz bandwidth can be doubled through the use two orthogonal
polarizations -- left-handed circular polarization and right-handed circular polarization. Further
study is needed, however, to see if the polarizations can provide enough separation to satisfy the
high signal-to-ratio requirement of 256-QAM. In addition, depolarization of the signal in the
presence of water vapor in the atmosphere can cause the two polarizations to interfere with one
another. An equalizer can be used to alleviate the effects of depolarization.
The cost for the high-rate system only includes telecom system on the spacecraft and comes up to
about $ 15 M. It does not include the necessary upgrades of the DSN sites such as new RF frontend
at 37 GHz, high rate baseband receiver, and equalizer. All ground station development cost
and operation expenses are kept separately in the ground systems cost estimate. The ARISE
spacecraft needs to orient itself to point the inflatable antenna at the observed objects. One
gimbaled antenna is needed to meet the 70% availability requirement. The gimbaling system will
allow for a half space view of the Earth (180 deg. 2 DOF capability).
There are many challenges for using 256-QAM as the 8 Gbps system. High order modulations
such as 256-QAM has thus far only been used in very stable communication channels like wire-line
systems. Atmospheric effects can make reliable transmission of 256-QAM difficult. The effect of
water vapor at 37 GHz can be significant especially in heavy rain at low elevation angles. Highly
linear power amplifiers at 37 GHz need to be developed. There are also proposed lunar missions
with whom ARISE is expected to share the 37-37.5 GHz spectrum. Although the overlapping of
telecom coverage areas on the Earth is not expected to be significant, ARISE will have to
coordinate with these missions to avoid mutual interference. A separate transmitter using only the
37.5-38 GHz spectrum can be added to provide downlink during overlaps, albeit at a lower data
rate.
We expect the advances in high data rate commercial RF systems will solve many of the problems
described above by 2004. The risk of the RF system can also be significantly reduced if the
required data requirement is lowered to 2 or 4 Gbps. One problem which deserves immediate
attention that the current stage of the design effort has not been able to address is the cross coupling
of the 1.2-m telecom antenna and the 25-m inflatable antenna. The downlink frequency of 37-38
GHz of the telecom system is very close to two of the ARISE observation bands at 43 GHz and 22
GHz. The transmit signal of the telecom system is many orders of magnitude larger than the
observed signal at 43 and 22 GHz and the transmit power spectrum of the telecom system may not
undergo sufficient attenuation at these 43 GHz and 22 GHz. Cross coupling of the transmit signal
to the 25-m antenna can contaminate the signal in the observed bands. Judicious placement of the
antenna and the use of absorbing material and other techniques should be investigated.
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Transmitter power 30.00 Watts
Transmitter power 44.77 dBm
Transmitter losses - 2 . 0 0 dB
Antenna gain 53.00 dBi
Antenna Efficiency - 2 . 2 2 dB
Pointing loss - 3 . 0 0 dB
EIRP 90.55 dBm
Distance 4.00E+04 km
Link Frequency 3.75E+10 Hz
Atmospheric attenuation - 5 . 0 0 dB
Space losses - 2 1 5 . 9 6 dB
Ground receiver parameters
Polarization losses - 1 . 0 0 dB
Receive antenna gain 80.00 dBi
Receiver cable/feeder losses - 2 . 0 0 dB
System Noise Temperature 80.00 K
Noise spectral density -179.57 dBm/Hz
Received power Summary
Received total power - 5 3 . 4 1 dBm
Received Pt/No 126.16 dB-Hz
Data Rate 4.00E+09 bps
Eb/No 30.14 dB
Eb/No Threshold (uncoded) 24.00 dB
Link Margin 6.14 dB
Table 6.4. Link budget of the 8 Gbps downlink. The link is consist of a RHCP and a LHCP
each at 4 Gbps.
64-QAM
Modulator
SSPA Data In 256-QAM PA
256-QAM
Modulator
PA
Data In
up to 4 Gbps
up to 4 Gbps
Shaping
Filter
Shaping
Filter
LHCP
RHCP
1.2 m
HGA
Figure 6.3. Block diagram of the 8 Gbps downlink transmitter
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6.5 Spacecraft thermal design
The thermal control system for the ARISE spacecraft consists of two specific elements: 1) the
cryocooler stage, and 2) the bus thermal control.
6.5.1 Cryocoolers stage
The cryocooler stage will consist of a three stage cooler system, which is required to provide the
science instruments with a 20 K temperature. The first stage will be a Stirling mechanical cooler,
while the second cooler stage will be a Sorption cooler. The first stage will operate between about
295 K to 60 K, while the Sorption stage will bring the science detector to 20 K. This system will
require 350 watts of electrical power, and will have a mass of about 90 kg.
6.5.2 Bus Thermal Control
The other spacecraft systems that affect the Thermal Control System are: the Power system,
because of the batteries, and Solar Array requirements; the Propulsion System, because of the
temperature requirements of the propellants; and the systems that require electronics components,
because of their temperature limits. Further the thermal design requires a knowledge of the
structure because its material (thermal conduction) and configuration (radiation) effect the thermal
exchange between spacecraft elements.
The TCS must control the temperature of the spacecraft elements within allowable limits for this
spacecraft, which has an electrical power level of about 2400 watts, has a cold zone, which must
be maintained at 20 K. The design uses standard passive thermal control elements, and will use
technology that is available at the technology cut off date. Multilayer Insulation (MLI) blankets will
control the thermal radiation between the spacecraft and space as well as between spacecraft
elements. Thermal surfaces will be used to control the thermal balance between the spacecraft and
the environment. Thermal conduction control will be used to maintain thermal gradients as
required. Also required are electric heaters and controllers for temperature sensitive elements such
as the batteries, and propulsion elements.
To maintain the science elements at 20 K, a two stage cryogenic cooler system is required, and will
consist of a Stirling cooler, and a Sorption cooler system. The design must incorporate thermal
isolation between the spacecraft bus elements and science stage, which will require thermal
conduction and radiation isolation. The thermal energy from the cryogenic coolers will be
transferred to thermal radiators with looped heat pipes that are mounted on the spacecraft bus. The
thermal radiators will be constructed from high performance composite material and will also
incorporate heat pipes.
The thermal control of the inflatable elements will use passive means, plus heaters, if necessary for
storage, deployment and rigidization. Several optional rigidization techniques are being evaluated,
one technique is cold rigidization. This technique requires that the inflatable elements be kept
below 225 K. An initial analysis shows, that with the correct external thermal surface, in this case
FEP-Aluminum or FEP-Silver, this temperature level can be achieved. To provide the uniformity
required, a simple 5 layer MLI blanket will be necessary. The inflatable elements must be kept
above the rigidization temperature during launch and prior to deployment, and this will be
accomplished with a MLI cover, and a small heater. The deployment must be accomplished rather
rapidly, as the inflatable elements will cool to 225 K between 3 to 20 minutes. Figure 6.5
summarizes thr thermal control system elements for ARISE.
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System Elements Mass (kg) Electrical Power (watts)
Bus Elements
Multilayer Insulation (MLI)
Thermal Conduction Control
Thermal Control Surfaces
Thermal Radiators
Thermal Louvers
Looped Heat Pipes
Electric Heaters/Thermostats
Instrumentation
Misc.
14.0
3.0
2.0
11.0
4.0
3.0
5.0
2.5
20.0
100 Avg.
Cryocooler Elements
cryocooler 90.0 350
TOTAL 154.5 450
Table 6.5: Thermal Control System Elements
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6.6 Spacecraft attitude control
The ACS system (in conjunction with propulsion) has to perform changes in velocity (delta-V). In
addition, it must determine and control spacecraft attitude and rate to allow science observations. It
must do this in the presence of various external and internal disturbances. To verify performance,
models must be built for both static and dynamic analysis. Below we discuss the requirements,
choice of components, cost, and analysis including modeling.
ACS Requirements
The ARISE pre-deployment requirements include, from the ACS standpoint, a 380 m/s delta-V
(approximately 34 minutes duration given a 450 N main engine). The science requirements are
described as follows:
- calibration requires a 2 degree slew in 60 seconds;
- a slew of 180 degrees in 60 minutes to 2 hours is desirable (slew rate 1 to 4 degrees in 60
seconds);
- keep the boresight to within +/- 30 degrees from the Sun;
- a quiescent phase during observing time of a duration from 2 to 20 minutes (depending on
the observation frequency);
- pointing accuracy during observation of 2 to 6 arcseconds.
The ARISE stability requirements, as a function of frequency, are shown in Table 6.6.
While maintaining these requirements the ACS must also counteract external disturbance torques
consisting of Earth's gravity gradient forces and moments, and solar pressure forces and moments.
In addition there are internal disturbance sources such as the ACS components themselves
(thrusters or reaction wheels) and other devices such as certain coolers (sorption coolers will be
quiet, but Stirling can be quite noisy).
Frequency [GHz] Motion [arcsec] Time Scale [sec] Stability [arcsec/sec]
5 50 350 0.042
8 29 350 0.024
22 11 150 0.020
43 5 60 0.028
86 3 15 0.025
Table 6.6: Stability Requirements
Components
To satisfy these stringent requirements and perform routine ACS operations, one set of reaction
wheels and two sets of thrusters with different thrusting capability are envisioned as the actuators.
One star tracker, one Sun sensor, one Inertial Reference Unit, and one GPS receiver are
envisioned as on-board attitude sensors. These may be redundant for reliability as desired. These
components are described in Table 6.7.
The ACS design is driven by the tight requirements and low structural frequencies of the antenna,
which dictates reaction wheels for fine pointing. These are sized by the torque and momentum
capability required for slewing and counteracting environmental torques.
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Vibration isolation components may be necessary, depending on the design of the cooling system,
and the results of more detailed dynamics simulations. These can either be passive, as the isolation
used for the Hubble Space Telescope reaction wheels, or active, as for STRV2.
These components can meet the accuracy requirements for the spacecraft bus itself. However, to
point the optical boresight to these same accuracy will require calibration of the alignment between
the optical axes and the bus axes. In addition, the stability of this calibration is an issue since it may
not be possible to calibrate during observations. Thermal variations and material aging may cause
significant perturbations requiring periodic re-calibration. This issue may require a closer
interaction between the RF and ACS subsystems. This may also require an active metrology
system for calibration during observations.
Component Type Mass [Kg] Power [W] Number
Reaction Wheels
Electronics
Teldix DR50 12
2
150/15
30/5
4
4
Star Tracker CT601 class 8 12 1
IRU HRG 5 22 1
Sun Sensor Electronics Head 0.5 0.5 1
0.9 N Thrusters 2.5 5 8
22 N Thrusters 2.5 5 8
ACS Computer
TOTAL 62 122
Table 6.7. ACS Hardware Components
Performance Analysis
Some calculations have been done to size the disturbance environment and quantify the pointing
problem. A finite element model of the ARISE spacecraft has been built in Matlab using the IMOS
software (Integrated Modeling of Optical Systems) . This has been refined by using NASTRAN
data for consistency with the structural design. The finite element model features all the structural
dynamic components of the spacecraft, with the exception of the bus and the subreflector, which
are assumed to be rigid. The solar panels and the subreflector boom are, however, modeled using
finite elements. See Figure 6.4 for the model. Therefore, we have beam elements for the support
struts and the hard truss, and membrane elements for the reflector and the canopy. The inflatable
torus, modeled as a circular ring, is connected to the reflector/canopy through a set of pretensioned
constant force springs. The membrane elements are linear, with no pretension. All material
properties are homogeneous and isotropic. The finite element model has 1876 degrees of freedom
(132 beams, 72 constant force springs, 396 membranes, 24 multipoint constrained degrees of
freedom, 472 massless degrees of freedom obtained through Guyan reduction), of which 1382 are
retained for the dynamic analysis. A model for the reaction wheels, including saturation at 0.2 Nm,
is also included in the structural model. The model also describes input forces and torques, such as
those derived from gravity gradient, solar pressure, thruster forces, and reaction wheel torques.
Static and dynamic deformation under open loop or closed loop control can be produced.
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Figure 6.4.
Disturbances
Given the inertia matrix, it is easy to determine the gravity gradient torque for arbitrary spacecraft
orientations. The worst case gravity gradient torque is less than 5.12E-3 Nm. IMOS also has the
capability of determining the gravity deformation forces that also result from the gradient; these are
less than 4.5E-4 N. The solar force direction in the spacecraft frame of reference has an angle with
the boresight (theta), and an angle of rotation around the boresight (alpha). The impact of the force
on the antenna can be determined and the forces and torques determined for various geometries.
See Figure 6.5.
Preliminary analyses show that the solar force is less that 3.7E-3 N, the solar torque is less than
0.05 Nm, the gravity gradient force is less than 4.5E-4 N, and the gravity gradient torque is less
than 5.12E-3 Nm. Of interest is the distance between the center of pressure and the center of
mass, equal to [-6;-7.5;-14.0] m. Also, see Figures 6.6, 6.7, 6.8, and 6.9 for the various dynamic
quantities of interest during the solar torque unloading maneuver.
Cooler disturbance data is not yet available, but an example of the possible magnitude of the
disturbance is given by work on STRV2. In that case, a 1 watt TI cryogenic cooler was used. It
produced forces of about 5N at various harmonics of the 55 Hz drive frequency. If the ARISE
cryocooler produces forces of this magnitude, it is very likely that it will have to be isolated at least
by a passive system similar to that used on Hubble.
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Figure 6.5
Figure 6.6
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Figure 6.7
Figure 6.8
Envelope
Components of the torque
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Figure 6.9
Momentum Management
After examining various options for reaction wheels, we have chosen the following wheel:
Teldix DR 50:
- torque: 0.3N
- max momentum: 300 Nms
- max wheel speed: 6000 Rpm
- power: 150/15/3 watt
- mass: 12 kg
- size: 0.15 m x 0.5 mD
Given the maximum torque, we can wait as long as 99 minutes before unloading the wheels,
which certainly is longer than required. The wheels will then be spinning at maximum speed (6000
RPM) and drawing maximum power. A better choice seems to be to unload if the wheels reach
about 1/4 of their momentum capability, which requires much less power. This also leaves a large
margin for observational flexibility. In addition, the reaction wheel disturbances are functions of
the square of the wheel speed, so minimizing the speed improves the pointing performance.
The power usage, considering all solar angles is given in table 6.6. The max power is the power
required for all 3 wheels just before unloading, minimized over all solar incidence angles. This
assumes we unload all wheels at once. The maximum average power is the power for all 3 wheels,
averaged over one cycle, taking the maximum over all solar incidence angles. When the antenna is
JPL Document # 16330
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pointing at 30 degrees from the sun (worst case requirement), it can stay at minimum torque
capability of the reaction wheels for a total of 31 minutes, before unloading. To be able to sustain
the solar torque disturbance, it can spin at 1.2 mrad/s for about 12 hours.
Table 6.8. Power for reaction wheels during observations.
option time before
unloading (min)
H at
unloading (Nms)
max power
(watt)
max average
power (watt)
A 99 300 314 180
B 26 80 117 81
The wheels can perform 2 degree slew in about 118 seconds and a 180 degree slew in about 19
minutes. Both of these are well within the requirements.
Unloading the solar torques can require a significant amount of hydrazine. This amount can be
reduced by using ion thrusters. This amount can also be reduced by rotating the whole spacecraft
about the boresight, but this maneuver would affect the power collection and telecommunications
subsystems, and to some extent even the science data gathering.
Dynamic Analysis and Control
While we do not want to use the thrusters during science observations, it will be necessary to
unload the reaction wheels periodically and so we want to quantify the disturbance this will cause.
Preliminary analysis of a 2 second firing of a pair of 0.9 N thrusters resulting in a couple about the
vertical axis of the spacecraft (z), shows that the maximum relative deformation at the joint between
the torus and a rigidizable struts is never exceeding 20 mm, and the residual vibration rapidly dies
out because of the high structural damping present in the inflatable structure (3% structural
damping). See Figures 6.10, 6.11 and 6.12.
Figure 6.13 shows the attitude control block diagram used in the simulations. Figure 6.14 and
Figure 6.15 show open loop simulation done with the Hubble reaction wheel model, at 500 rpm
and 2000 rpm, respectively. What is shown is the angle due to deformation at the torus-strut
attachment point when the wheels are operating. Based on test data, the wheels produce
disturbance forces and moments due to imbalance, motor cogging, and ripple. For the 2000 rpm
case we can still meet the requirements. This indicates that isolation of the wheels may not be
needed. Note that we may obtain additional margin by keeping the wheels at a lower speed by
unloading more often, which is quite possible as the observation times for the radio frequencies
requiring the highest precision are only about 2-3 minutes. If we unload every 10 minutes for the
maximum disturbance torque, then the peak wheel speed is only about 600 rpm.
Closed loop analysis was made of a 2 degree slew in 200 seconds maneuver about the x-axis of the
spacecraft, using reaction wheels. The results show this to be a very benign maneuver. Figure
6.16 shows the spacecraft slew angle, Figure 6.17 the reaction wheel torque profile, and Figure
6.18 the displacement at the strut-torus attachment.
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Figure 6.10
Figure 6.11
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Figure 6.12
Figure 6.13
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Figure 6.14
Figure 6.15
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Figure 6.16
Figure 6.17
r
a
d
JPL Document # 16330
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Figure 6.18
6.7 Structures and mechanisms
As discussed in the spacecraft configuration section, the spacecraft bus has an octogonal shape 2 m
long and 1.3 m wide. It was not attempted in this study to define and design the bus material and
thickness. A mass of 10% of the spacecraft dry mass minus propulsion and inflatable antenna
subsystem masses was allocated for the spacecraft bus structure, which rounds up to about 71 kg.
Masses for the mechanisms, such as subreflector truss and deployment, solar array gimbals,
telecom antenna boom and deployment, were estimated. Cables and connectors were taken as
7.5% of the spacecraft dry mass minus propulsion and inflatable antenna subsystem masses. An
allocation of 10 kg was also made for additional radiation shielding of sensitive parts of the
spacecraft (C&DH shielding was bookkept separately).
6.8 Power subsystem
The power system has three major parts. The solar array provides power during sunlit periods.
The battery provides power during eclipses, supplements the solar array during peak power
periods, and provides power during the immediate postlaunch period, before the solar arrays are
deployed. The PMAD system provides power management and distribution. It includes the peak
power tracker; distribution, regulation and control electronics; and pyro.
Calculations of the estimated solar array area and mass were based on spacecraft requirements of
2270 W EOL, which includes 30% contingency. Until further details are available on the spacecraft
power profile, it was assumed that the solar array would handle all power needs during sunlit
periods. Adding an estimated 125 W to recharge the Li-ion secondary battery, the overall array
JPL Document # 16330
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sizing assumed a net 2400 W EOL requirement. The results for six of the leading cell candidates
are detailed in Table 6.9. These mass and area numbers include the cells; thin coverglass (3 mil),
with the exception of the copper indium diselenide (CIS) cells which do not include coverglass;
wiring, terminals, connectors, and substrates. As is customary, they do not include additional
contingency (this is carried at the system level), nor do they include the support structure
(connection to the spacecraft), deployment, drive or housing. Overall, the most reasonable
compromise between area, mass, cost and availability was projected to be the inflatable array
(ITSAT type) using high-efficiency Si cells at a specific power of 86 W/kg BOL. The array area
would be 16.3 m2
and the array mass would be 32.8 kg.
Area (m2) Mass (kg)
GaAs 14.9 46.3
2-junction (GaInP/GaAs) 13.3 41.1
3-junction (GaInP/GaAs/Ge) 11.1 32.6
CIS (LMA est.) 36.9 34.3
CIS (L’Garde est.) 36.9 25.2
High efficiency Si 16.3 32.8
Table 6.9: Calculated Solar Array Area and Mass for 2400 W EOL
GaAs = gallium arsenide on Ge substrates
GaInP/GaAs = two junction cell on Ge substrate
GaInP/GaAs/Ge = three junction cell on Ge substrate; includes active Ge junction
CIS = copper indium diselenide
Calculations of the estimated secondary battery mass and volume assumed that the battery would
be used only during eclipses (460 W for 45 min, or 345 Whr). The primary battery requirements
cover a 8.8 hr period immediately postlaunch (4930 W-hr). Until further details are available on the
spacecraft power profile, it was assumed that the solar array would handle all active power needs
during sunlit periods. It was assumed that neither science data collection nor telecom would not
occur during eclipse. It was also assumed that the mission lifetime would be limited to about 3
years, in order that a Li-ion battery could handle the required number of cycles. The numbers do
not include battery mass or battery volume contingency, which would be carried at the system
level. The required 25 Ahr Li-ion secondary battery would have a mass of 7.3 kg and a volume of
6 liters. It is evident that a large mass and volume penalty would result if a Ni-based battery were
to be substituted.
Calculations of the estimated PMAD mass were based on extrapolations from the Phase A Light
SAR calculations performed at JPL in 1996. It was assumed that the ARISE EOL solar array
power would be 2400 W and the secondary battery capacity would be 25 Ahr. The calculated mass
of the peak power tracker would then be 13.9 kg, and the mass of the distribution, regulation and
control electronics would be 66.7 kg, for a total PMAD mass of 80.6 kg. The corresponding EOL
PMAD specific power would be 30 W/kg. This corresponds favorably to the 30 W/kg anticipated
for the JPL X-2000 PMAD second delivery. The results are detailed in Table 6.10.
Several key technology challenges for the power system were identified. First, the advanced solar
array technologies (multijunctions and CIS) must be scaled up without appreciable efficiency loss
if they are to compete effectively with Si and GaAs. Second, deployment mechanisms for
ultralightweight solar arrays need to be flight qualified. Third, Li-ion secondary batteries need a
flight demonstration; they also need to be demonstrated in large sizes (over 20 Ahr). Fourth, the
PMAD mass can only be reduced if the projected parameters of the X-2000 3rd delivery
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(approximately 200 W/kg) can be demonstrated and scaled up. The X-2000 3rd delivery is planned
for a very small (10 W) power system.
Extrapolated from LightSAR case:
LightSAR ARISE
Solar array power (EOL) 782 W 2400 W
Battery capacity 44 Ahr (Ni) 25 Ahr (Li)
Peak power tracker 6.4 kg 13.9 kg
Dist, Reg & Cntrl Electronics* 15.0 kg 66.7 kg
Total PMAD mass 21.4 kg 80.6 kg
PMAD specific power (EOL) 36.5 W/kg** 30 W/kg***
Table 6.10: Calculated PMAD Mass
*Estimate based on EOL array power corrected for environmental degradation only (EOL/0.85)
**LightSAR assumed a very bare-bones system
***30 W/kg is approximate value for X-2000 PMAD 2nd delivery
In summary: Size estimates, including mass and area, have been generated for the ARISE solar
array. Size estimates, including mass and volume, have been generated for the ARISE battery. A
ROM estimate of power electronics mass and specific power has been calculated. Several key
needs were identified. First, more data on the planned orbit and eclipses are needed in order to
refine the power system sizing. Second, more data on the spacecraft power profile vs. time are
needed in order to determine the proper role of the battery in supplementing the solar array. Third,
the calculations assume a moderately benign radiation environment, which may not actually be the
case in the planned ARISE orbit; more data is needed on the radiation environment in order to
properly size the power system, particularly the solar array which is relatively difficult to shield.
Planned near-term activities include: updating the power system design in accord with evolving
system requirements; continuing to reduce the power system mass; refining the PMAD mass and
cost estimates; and establishing a power system design and fabrication schedule.
Backup data are available in the tables in the Appendix G (Tables A through G).
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6.9 Propulsion subsystem
The propulsion module is a bipropellant dual-mode propulsion system that is used to perform a
~380 m/s periapse raise, reaction control during the periapse raise, and attitude control for the
duration of the mission.
The bipropellant dual-mode propulsion system uses nitrogen tetroxide (NTO) and hydrazine
(N2H4) as the oxidizer and fuel, respectively. The periapse raise is performed using a 445 N Royal
Ordinance LEROS-1c main engine that is qualified for these propellants exclusively. Two titanium
tanks (one for the oxidizer and one for the fuel) are used to store the propellant. The oxidizer and
fuel tanks are pressurized via separate high-pressure helium feed systems (two pressurant tanks).
The separate feed systems eliminates any possibility of propellant migration. Eight 22 N and eight
0.9 N monopropellant (hydrazine) thrusters are assumed for thrust vector and roll control.
Conventional technology components are assumed.
This design has two possibilities for combining the inflation system with the propulsion system.
One option is to have an NTO inflation system feeding off a line downstream of the oxidizer tank.
Liquid NTO decomposes into gaseous N2 and O2 through a two-step reaction. Another option is to
have an N2H4 inflation system feeding off a line downstream of the fuel tank (feeding off the RCS
system). Liquid N2H4 decomposes into gaseous NH3, N2, and H2 through a two-step reaction.
Since there are separate pressurization feed systems for both the oxidizer and the fuel, both tanks
remain pressurized for the entire mission. No pyrotechnic firings are necessary after the periapse
raise.
P
NTO
GHe T
T
P
P
N2H4
GHe T
T
P T
T
T
T
T
(on each catalyst
bed heater)
(on each thruster)
T T T
T
T
(on each catalyst
bed heater)
(on each thruster)
22 N Roll Cntr. 22 N TVC
0.9 N ACS
0.9 N Roll Cntr.
450 N Main Engine
ARISE S/C Propulsion System: Option #3
P
To inflation system
(alternative B)
P
T
GHe
P
T
GHe
To inflation system
(alternative A)
Legend
Latch Valve
Service Valve
Filter
T Temperature Transducer
P Pressure Transducer
Gas Regulator
Orifice
Test Port
Pyrotechnic Valve
(normally closed)
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7. Ground systems and mission operations
The FY’98 work focused on the space segment of ARISE. The issues and design considerations
associated with the ground segment will be studied in more detail in FY’99. This section
summarizes the current understanding of the ground system and mission operations.
Operations and data handling scenario
The ARISE Mission carries out observations on an approximately 70% duty cycle. During
observations instrument data are immediately sent to the ground, so the spacecraft must be tracked
during all observation periods. Instrument data loss of up to 20% is tolerable during these tracking
periods. Instrument data is transmitted at 8 Gbps over a Ka-Band link to a set of dedicated ground
terminals. It is recorded at 8 Gbps on tapes which are then shipped to a VLBI data processing
center. This instrument data flow is shown on the lower part of Figure 7.1. The upper part of the
figure shows the downlink and uplink flows for engineering telemetry and spacecraft commands.
Engineering data is recorded on-board the spacecraft and is played back once a week over a DSN
34 meter tracking pass. During this pass the commands to control the next week’s worth of
observations are transmitted to the spacecraft. These commands are the result of the planning of
science and engineering activities needed to achieve the mission goals. A coordinated set of
observation plans are sent to radio telescopes on earth to direct the collection of concurrent
observations to be correlated with the observations conducted from the spacecraft. Details of the
ground system components needed to operate the mission are discussed in the next section.
Space
Radio
Telescope
S/C
Data
System
S/C Ka
-Band
Comm
System
Dedicated
Ground
Terminal
Network
S/C RF
Comm
System
DSN
ARISE
Operations
Center
Cmd./Tlm.
Processing
Services
Science
Data
Recording
VLBI
Data
Processing
Center
ARISE
Science
Planning
Cmds.
Instr.
Data
Ref.
Frequency
Instr.
Data
Playback
Telemetry
Cmds.
Uplink
Signal
Downlink
Signal
Uplink
Signal
Downlink
Signal
Playback
Telemetry
Cmds.
Playback
Telemetry
Cmds.
Instr.
Data
Space Instr. Data
Earth
Radio
Telescopes
Science
Data
Recording
Earth
Instr.
Data
Earth
Instr.
Data
Observation Plans
Observation Plans
Science
Data
Products
Figure 7.1: End-to-End Data Flow
Ground system design
Figure 7.2 illustrates the ground system design for ARISE. The ARISE ground system is above
and to the right of the dashed line in the figure. The VLBI science observations are planned to be
executed by concurrent use of the spacecraft-based radio science instrument and a set of earthbased
radio science instruments. The coordinated observation plans are issued from the science
planning function shown on the far right hand side of Figure 7.2. Observation plans for the
spacecraft are translated at the ARISE Operations Center into a set of commands to be uplinked
through the DSN’s multimission command processing service once a week. During this once a
week DSN track the spacecraft plays back the last week’s worth of recorded engineering telemetry
which is delivered by the DSN’s multimission telemetry processing service. The ARISE
Operations Center analyzes the engineering data to assess the performance of the spacecraft and
generates any ancillary data about the spacecraft status needed to support VLBI data processing.
JPL Document # 16330
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ARISE will use a dedicated ground terminal network to handle the instrument data downlink. The
network will consist of 3 terminals distributed at sites around the world so as to make coverage
almost continually available to the orbiting spacecraft. The terminals will be remotely controlled
from the ARISE Operations Center, except for maintenance and tape loading, unloading, and
shipping operations. The instrument data from space will be recorded on 2 to 4 tapes for each
observation. The tapes will be played back at the VLBI Data Processing Center, along with tapes
containing radio science instrument data from concurrent earth-based observations. The VLBI Data
Processing Center will perform correlation, fringe fitting, image processing, and data archiving.
Dedicated
Ground
Terminal
Network
DSN
ARISE
Operations
Center
Cmd./Tlm.
Processing
Services
Science
Data
Recording
VLBI
Data
Processing
Center
ARISE
Science
Planning
Uplink
Signal
Downlink
Signal
Uplink
Signal
Downlink
Signal
Playback
Telemetry
Cmds.
Playback
Telemetry
Cmds., Tracking
Requests,
Ephemeris
Instr.
Data
Space Instr. Data
Earth
Radio
Telescopes
Science
Data
Recording
Science
Data
Products
Observation
Plans
Observation Plans
S/C
Instr.
Data
Earth Instr. Data
Tracking Schedule, Ephemeris
Tracking Status
Ancillary
Data
Figure 7.2: Ground System Functional Block Diagram
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8. Costs
The ARISE cost used JPL’s Team X cost estimation tools. Comparison of the ARISE cost profile
with other current flight mission profiles were done to validate its applicability. The Team X model
build for ARISE uses quasi grass roots estimates for the spacecraft subsystems, mission
operations, science team and launch vehicle. It uses historical estimate models for other mission
components. These costs assume a 20% reserves on phase A-D and 10% on phase E. They also
assume a phase A duration of 18 months, phase B duration of 18 months, phase C/D of 48
months, and phase E of 39 months. Redundancy is typical and the costs are estimated in FY’98
dollars. Table 8.1 summarizes the ARISE cost breakdown. More details can be found in Appendix
H. The known subsystem grass roots cost estimates are given below.
Phase A-D Phase E
Project Management 12.3 0.7
Science 2.8 1.3
Project & Mission Engineering 5.4
Payload (science instruments) 27.4
Spacecraft
System management
System Engineering
Inflatables
ACS
C&DH
Telecommunications
Power
Propulsion
Structure/thermal/cabling
Thermal control
Cryocoolers
Software
LV adapter
144.1
1.7
2.5
6.0
32.0
6.0
25.3
11.0
12.0
24.7
2.9
11.0
7.0
2.0
ATLO 20.1
Mission Operations 35.0 17.0
Launch Vehicle 60.0
Reserves 49.4 (@ 20%) 2.0 (@ 10%)
Total 356.6 21.0
Table 8.1: ARISE cost breakdown in FY’98 dollars
Ground system development cost and operations cost
The costs for the ARISE ground system development and ARISE mission operations are
summarized below. Each of the components of the ARISE ground system pictured in Figure 7.2
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are listed in the left hand column (see Table 8.2). The next 3 columns show the cost in FY 98
dollars for technology development, ground system development (Phase C/D) and mission
operations (Phase E). There is no need to provide advanced technology development funds for
any of the elements of the ARISE ground system. All technologies will be fully developed and
ready for use by the beginning of Phase C/D. The Phase C/D development costs are estimated to
total approximately $35M. The Phase E operations costs are estimated to total approximately
$17M for 3 years of operations.
Component Technology
Development
Costs ($M)
ARISE Development
Costs ($M)
Operations Costs (3
Years, $M)
DSN Tracking 0 0 0.8
Cmd./Tlm.
Processing Services
0 0.5 0.2
ARISE Operations
Center
0 3 7.5 (17 operators)
Dedicated Ground
Terminal Network
0 6 (1 terminal with control
system at each of 3 sites =
3, 1 high rate receiver and
equalizer at each site = 3)
4.5 (6 operators, tapes and
shipping for 3300 tapes [2 -
4 tapes per observation])
Science Data
Recording
0 9 (1 decoder at each site =
3, 2 recorders at each site =
6)
Included in Optical Ground
Terminal Network Costs
VLBI Data
Processing Center
0 16 (1 VLBI correlator and
image processor = 6, 10
recorders = 10)
3 (5 operators)
ARISE Science
Planning
0 0.5 1
Totals 0 35 17
Table 8.2: Estimated Development and Operations Costs
Telecom system cost
System Costs ($ k)
RF X-Band
Transponder 400
SSPA 400
Oscillator 500
Antenna 300
Diplexor 300
Cables and connectors 100
Power supply 600
Testing and support 300
RF Ka-band
Modulator + shaping filter 800
Power Amplifier 3,750
Antenna 4,000
Misc. electronics 2,000
Cables and connectors 200
Power supply 600
Testing and support 300
DSN
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Total 13,750
Thermal control system cost
Workforce * (11 WY x $ 150 K) 1,650 K
Multilayer Insulation Mat/Fab/Instl (14 kg) 700 K
Thermal Conduction Control 100 K
Thermal Control Surfaces 200 K
Thermal Radiator ( includes heat pipes ) 400 K
Thermal Louvers (4 units) 800 K
Loop Heat Pipes 1,000 K
Electric Heaters/Thermostats 200 K
Instrumentation 100 K
cryocooler 3,200 K
*This costing does not utilize the DNP process
Table 8.3: Estimated Thermal Control Subsystem Cost
Attitute control system cost
ACS Subsystem cost information is shown in Table 6.14.
ACS Subsystem Description Cost [K$]
System Engineering 660
Controls & Analysis 770
Software 1288
I & T 7007
GSE 2908
H/W Engineering 2789
Flight Hardware 16107
TOTAL ACS COST 31529
Table 8.4: ACS Subsystem Cost
Power system cost
A rough order of magnitude (ROM) cost estimate for the power system was generated. The details
are provided in Table 8.5. The estimated cost of the solar array was $3.5 M including the array
subcontract, JPL engineering support and all burdens. The estimated cost of the battery was $1.3
M including the battery subcontract, JPL engineering support and all burdens. The PMAD cost
was not yet defined since the details of the PMAD system were not yet available; however, it was
noted that the LightSAR PMAD system, for a smaller (782 W EOL) power system, was estimated
to cost $5.9 M including labor, engineering support and burdens.
Table 8.5: Power System Cost Estimate
• Solar array
– GaAs $2K/W x 1700 W = $3.4 M
– Si or CIS $1K/W x 1700 W = $1.7 M
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– High effic Si $1.5 K/W x 1700 W = $2.55 M
• CIS potentially even less expensive but not yet demonstrated
– Multijunction $2.5 K/W x 1700 W = $4.2 M
– Add JPL engineering support ($200 K/yr) + proc/general burdens
– Assume high efficiency Si baseline
• ((2.55*1.027) + (0.75*0.200) + (2*0.200) + (0.75*0.200))*1.058 = $3.5 M
• Battery
– Li-ion batteries = $170 K apiece x 3 batteries = $0.51 M (2005 est.)
• need flight, spare and qualification units
– Add JPL engineering support ($200 K/yr) + proc /general burdens
• ((0.51*1.027) + (0.75*0.200) + (2*0.200) + (0.75*0.200))*1.058 = $1.3 M
• PMAD
– LightSAR $5.9 M incl. labor + engineering support + burdens
– ARISE being calculated
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APPENDICES
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APPENDIX A
ARISE Mass Budget
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APPENDIX B
ARISE Power Budget
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APPENDIX C
ARISE Structures and Thermal Analysis
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APPENDIX D
SEP System
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APPENDIX E
ARISE radiation environment
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APPENDIX F
ARISE ESD environment
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APPENDIX G
Power Subsystem
ARISE Power System Appendix : Backup and Supporting Data
Carol Lewis, Sal DiStefano, Gene Wester
Prepared 3/23/98; Data as of 1/30/98
Table A. Solar Array Assumptions
Cell effic (BOL) W/m2 (BOL) W/kg (BOL)
GaAs 19% 199 68
GaInP/GaAs 21.5% 225 77
GaInP/GaAs/Ge 24.3 % 255 87
CIS (LMA est.) 10% 93 100
CIS (L’Garde est.) 10% 93 136
High effic. Si (2005 proj.) 19% 199 86
• BOL numbers for 1-sun AM0, 28oC (baseline)
• Assume actual array operating temperature of 85oC
• 10% CIS cells avail est. 1999
– LMA est. originally provided 2/97
– L’Garde est. w/Al rigidization originally provided 3/97
Table B. Assumptions for the Solar Array Recharging the Battery
• Portion of solar array power is needed to recharge battery
• Assume battery charge efficiency = 0.79
• Assume battery discharge (energy) efficiency = 0.95
• Assume 3.5 hr available to recharge battery per orbit
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– Battery prefers charge rate of at least 0.1 C
– 200 W of solar array required to do this
– If recharge time is shorter, more array area/mass required
Table C. Estimated Degradation Factors for Solar Array
• For 3 - 5 yr mission (BOL vs. EOL) -Assumptions
– Temperature coefficient factor
0.873 for GaAs, 0.868 for III-V multijunctions, 0.715 for Si or CIS
– Cell packing factor 0.85 (15% of array area not covered with cells)
– Radiation degradation 0.85
– Temperature cycling 0.98
– Fabrication losses 0.98
– Micrometeorites 0.98
– Wiring/diode 0.96
– IR losses 0.98
– UV degradation 0.98
– Offpointing 1.00
Table D. Possible Radiation Environment
• Assume ARISE orbit 5,000 - 40,000 km altitude
• Radiation data from GaAs Solar Cell Radiation Handbook (B. Anspaugh)
• AP8 proton model
– high energy peak at L = 1.5 Re (0.5 Re from surface = 3186 km altitude)
– intermediate energies peak at L = 2 Re (1.0 Re from surface = 6371 km altitude)
• AE8 electron model
– inner zone L = 1.2 - 2.8 Re (0.2 - 1.8 Re from surface = 1274 - 11468 km altitude)
• peak at L = 1.4 Re (0.4 Re from surface = 2548 km altitude)
– outer zone L = 3-11 Re (2-10 Re from surface = 19113 - 63711 km altitude)
• peak at L = 4-5 Re (3-4 Re from surface = 19113 - 25484 km altitude)
– L = distance from center of Earth; R = 6371 km
Table E: Battery Assumption s
• Assume 30 min eclipse/orbit but during only 3 consecutive months of year
– 3 yr mission = 2417 eclipses (battery cycles)
– 5 yr mission = 4028 eclipses (battery cycles)
• 3 yr probably OK for Li-ion (nominally up to 2000 cycles at 50% DOD)
• 5 yr calculated for both Li-ion and Ni-based batteries
• Looked at 3 types of Ni-based batteries which can withstand many cycles at 35% DOD
– Common pressure vessel (CPV) NiH2 - available now
– Single pressure vessel (SPV) NiH2 - should be available near-term
– NiCd - available now
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Table F: Assumed Battery Baselines at BOL
Whr/kg Whr/liter Max DOD for max cycles
Li-ion 100 120 50%
CPV NiH2 35 25 35%
SPV NiH2 53 68 35%
NiCd 25 35 35%
Table G: Power Electronics Assumptions
• In general power electronics (PMAD) includes peak power tracker, and
distribution/regulation/control electronics
• More specifically includes DC/DC converters for each load, bus limiter, power control,
power distribution network, bench test equipment (BTE), ground support equipment
(GSE) and pyro.
• Flight hardware does not include BTE and GSE.
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APPENDIX H
ARISE Cost Estimates
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APPENDIX I
ARISE Team