Dec 3, 1990 - Management Proposal. 8.1 . ...... system will then become the backup communications link between the TV and the ...... RECOVERY: CORR.
NASW-4435
/// iii
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_NAL O_
A
_PACE _:EMOVAL
_YETEM
Written in Response RFP #ASE274L
Submitted Dr. The
George
University
Department and
To: W.
of
University
of
By: Inc. Texas
at Austin
of Aerospace Engineering Engineering Mechanics
December (_AgA-L_,-!_9::_7:_)
at Austin
of Aerospace Engineering Engineering Mechanics
SPECS,
Department and
Botbyl
Texas
Presented
The
to
FIN,_L
3,
1990
_.)73IGN
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_,PACt. CSCL
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D_S_GH OF
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_EMOVAL
SYSTEM
INC.
Presented SPECS, Erika Don
Carlson,
Chambers,
Andrew Richard The
December
Steve
Casali Geissler
Manfred
John
University
Inc.
Garner
Lalich,
Mach,
By:
Parry, of
Foley
Texas
3,
Leipold
at
1990
Weems Austin
Executive
Overview
Overview In Texas
the
at
and a
small
can
effectively
The
research
debris
stage
large
there
to
debris,
that Our
is
current
to
data
the
is
uploaded
the
vehicle
netting
captures
it.
returns
to
continues netting a
Earth
for nets.
taken the take for of
back ability
the
the
into
orbit
to
a 30
degree transfer
the for
capture
50 six
inclination
the
a
months
a
and
the
change
on
system the
outgoing
all
launched
the
to
netting
the modules
by
the
are
taken
and
repacking
they the
One is
vehicle
returns
mission,
debris.
and and
Once
refurbished,
typical
Next,
netting
new
refueling,
orbital
This
debris,
vehicle
debris, In
to
module
captured
near and
cloud.
the
are
netting
based
to
modules
of
is
a
inclination
area.
modules
and
orbit.
expended
reuse.
remove
size.
proceeds
with
are
in
parking
netting target
resupplied
modules
cm
tracking
the
pieces
been
debris
it
transfer
new
little
system
nets,
new the
pose
have
targeted
available
is
The and
netting
vehicle.
and altitude
the it
load.
Once
approximately the
where
or
lower
in
expended,
of
breakup
for
large
and
50
based
vehicle
debris
a
around.
degrees
ground
the
spacecraft.
with
vehicle
The
28.5
tracks
vehicle
ground
removal
a
deployed,
more
shuttle
from
at
the
capture
to
transfer
at
shielding
at SPECS, Inc. have medium sized orbital
debris.
a satellite transfer
to
has
uncovered
because
debris
1 cm
a
uses
expending
orbit
from
size
spacecraft
environment
located
system
is
are
space
shuttle
ranging
problem
research
maneuvered
the engineers removal of
medium
transfer
station's
from
designs
debris
capture
modules
space
the
the
current
the
the
After
to
many
of
to
with
large
and
station
location
the
space
and
collision
tracked
debris debris
damaging
though
of
The
spacecraft
from
incorporates
The
rendezvous
particles
analysis, on the
space
the
to
University
orbital
Our
be
design
altitude.
determine
threat
the
the
satellites
little
pieces
capture
to
(LEO).
from
those
operational km
are
From this concentrate
risk
even
the
system.
orbit
spacecraft,
can
particles
proposed. decided
the
of
studying
Earth
that,
the
they
debris
vehicle
showed
of
Inc.
removal
these
destroy
because
task
which
posed
also
SPECS,
debris
low
prevent
Additionally,
400
at in
1990,
the a
particles
could
danger
an
accepted
substantial
that
of
designing
reached
become
semester
Austin
problem has
fall
to of
then
system
mission
designed and
are
back
has will
to incoming
allow trips
Transfer
Vehicle
The transfer vehicle is the part of the debris removal system that moves the nets, netting vehicle, and netting modules close to the debris that is targeted for capture. A basic layout of the vehicle is shown in the following diagram.
Figure The
transfer
1
Transfer
vehicle
is
Vehicle
capable
of
30
Layout
degrees
of inclination
change on both legs of the trajectory. inclination change without massive vehicle uses ion engines for thrust. reduced to 10% of the amount that
To accomplish the large amounts of fuel, the transfer This allows the fuel amount to be would be used if chemical engines
were
used.
of power
require, vehicle in the
the transfer vehicle uses 2 high efficiency solar arrays. The also has batteries that will provide power while the vehicle is shadow of the Earth. The transfer vehicle weighs approximately 8,000 kg. When it
To
provide
the
35
kW
that
the
10
ion
engines
is fully loaded with the netting modules, propulsion module, and fuel, the transfer vehicle weighs 30,000 kg. Once the netting vehicle has captured mass of weight debris.
the debris the transfer is
due
Control
and returned to vehicle is about
to
the
fuel
that
of
the
transfer
is
spent
vehicle
the transfer 21,000 kg. during is provide
vehicle, the total This reduction in
the by
capture control
of the moment
gyroscopes. The gyros will perform the fine attitude adjustments required as the vehicle rendezvous with the debris. For large maneuvers and momentum dumping, the vehicle also includes RCS thrusts similar to those used by the space shuttle. onboard
Navigation of calculations
the transfer and data
vehicle is done by a combination from ground. Initially, the transfer
ii
of
vehicle receives data about the location of the debris and its location from external sources. From the data, the vehicle plots an intercept course. The vehicle proceeds along its trajectory and modifies it as new data is received about the location of the vehicle with respect to the debris. The transfer vehicle receives this data from the command center located on Earth via a Ku-Band communications link through the TDRSS satellite. The transfer vehicle relays any commands to the netting vehicle with a V-Band communications system. Netting Vehicle The netting vehicle is responsible for gathering the debris and returning it to the transfer vehicle. The layout of the netting vehicle and the modules is shown in the following diagram.
Figure
2
Netting
Vehicle
Layout
Once in the debris orbit, the netting vehicle infrared (IR) tracking system to locate and target
uses its onboard a piece of debris.
Once the debris is targeted, the netting vehicle does a Hohmann transfer into a slightly different orbit. This allows the netting vehicle to close in on the debris piece. As the vehicle closes in on the debris to a distance of about 25 km, the tracking switches to a LADAR (LAser Detection and Ranging) system. The LADAR system provides more accurate ranging and location information to the netting vehicle as it approaches the debris. When the debris is within about 20 m of the debris, the vehicle will fire a net, capture the debris, and reel the net
back into the netting module. The netting vehicle will be controlled by ground or elsewhere with teleoperated controls. This will prevent the netting vehicle from having to have extensive artificial intelligence. The communication is relayed to and from the netting vehicle using VBand link from the transfer vehicle through TDRSS. To provide the attitude adjustments, the vehicle conjunction with RCS thrusters.
will use control The vehicle will
iii
moment also use
gyros
in
Hydrazine/Nitros Oxide fueled engines to provide the large orbital changes as the vehicle chases the debris. Power is provided by surface mounted solar arrays. The arrays were surface mounted so that the area of the craft wasn't increased by the arrays. This is important because the smaller our craft, the less the chance of a collision with a debris particle. The array is also oversized by 25% to compensate for degradation due to debris impacts. The total mass of the netting vehicle after it leaves the transfer vehicle is 8076.5 kg. Upon gathering all the debris and returning to the transfer vehicle, the mass is reduced to 5183 kg. This reduction in mass is caused by expending the fuel.
iv
Table Executive Table of Table Table
of
Contents i
Overview ............. Contents ..................
V
viii
of Figures ..................... of Tables ......................
1.0 ............................................... 1.1 ................................... 1.2 ................................... 1.2.1 .................... 1.2.2 .................... 1.2.3 .................... 1.2.4 .................... 1.2.5 .................... 1.3 ...................................
Project Project Defining
Overview Objective and Scope the Debris Environment
Types of Debris Debris Location Sizes of Orbital Targeted Dynamics General
Debris
Debris Environment of Satellite Breakup Project Requirements
1.4 ................................... 2.0 ............................................... 2.1 ................................... 2.2 ................................... 2.4 ................................... 3.0 ............................................... 3.1 ................................... 3.1.1 ....................
Assumptions Design Approach Design Options Primary Design Design Philosophy System Concept Debris Removal Transfer Vehicle
3.1.2 .................... 4.0 ...............................................
Netting Mission
4.1 ................................... 4.2 ................................... 4.2.1 .................... 4.2.2 ....................
System
Vehicle Scenario
Mission Options Final Mission Scenario Active DRS Launch Orbital Transfer of the
4.2.3 ....................
Rendezvous
4.2.4 .................... 4.2.5 ....................
Netting Further
4.2.6 .................... 4.3 ...................................
Resupply Discussion
5.0 ............................................... 5.1 ................................... 5.1.1 ....................
Subsystem Propulsion Transfer
and
TV
Debris
Capture
Vehicle Resupply Orbit Changes Base of
on SSF Alternative
Missions
Design Vehicle
5.1.2 .................... 5.2 ................................... 5.2.1 ....................
Netting Power Transfer
Vehicle
5.2.2 .................... 5.3 ...................................
Netting Thermal
Vehicle Power Subsystem
Vehicle
Propulsion Power
Supply Supply
ix 1 2 3 3 4 6 9 9 10 11 11 12 12 13 14 14 15 17 19 19 20 20 22 22 23 23 23 24 26 26 26 32 34 34 37 38
5.3.1....................Transfer Vehicle 5.3.2....................Netting Vehicle 5.4 ................................... Communications 5.4.1....................Subsystem Requirements 5.4.2....................Design Approach 5.4.3....................Subsystem Design 5.4.4....................Netting Vehicle 5.4.4.1 ....V-Band Network 5.4.4.2 ....S-Band Network 5.4.5....................Transfer Vehicle 5.4.5.1 ....Ku-Band Network 5.4.5.2 ....V-Band Network 5.4.5.3 ....S-Band Network 5.5 ................................... Data Processing 5.6 ................................... Tracking and Detection Subsystem 5.7 ................................... Guidance, Navigation, and Control 5.7.1....................Guidance and Navigation 5.7.2....................Vehicle Control 5.8 ................................... Netting Subsystem 5.9 ................................... Structural Materials 6.0 ............................................... System Integration 6.1 ................................... Debris Removal System 6.2 ................................... Debris Retrieval 6.3 ................................... Docking 7.0 ............................................... Debris Prevention Concepts 7.1 ................................... Self Disposal of Spacecraft 7.1.1....................Drag Devices 7.1.2....................Solar Sails 7.1.3....................Deorbit Engine 7.1.4....................Additional Fuel 7.2 ................................... Subsystem Redesign 7.2.1....................Rocket Redesign 7.2.2....................Seperation Mechanism Redesign 7.2.3....................Use of Reusable Hardeware 7.2.4....................Improved Shielding 7.2.5....................Redesign of Protective Coating 8.0 ............................................... Management Proposal 8.1 ................................... Management Structure 8.2 ................................... Subgroup Responsibilities 8.3 ................................... Task Development 8.4 ................................... Workload Considerations 9.0 ............................................... Cost Proposal 9.1 ................................... Personnel Cost Estimate vi
38 38 39 39 41 43 44 45 47 49 5O 51 54 55 56 60 60 63 67 69 76 76 76 83 84 85 85 85 86 86 87 87 87 88 88 88 89 89 91 91 92 96 96
9.2 ................................... Material and Hardware Costs 10.0............................................. References Appendix A Spacecraft Anomaly Reports Appendix B Other Design Options Considered Appendix C Calculation of Perturbative Accelerations Appendix D Himawari 1 Rocket Booster Explosion Appendix E Fuel Calculation Program Listing
vii
97 98 104 109 112 113 115
Table
of
Figures
Figure Figure Figure Figure
1.1 Tracked Debris Separated into Groups ................................. 1.2 Global Outlook of the Debris Problem ................................... 1.3 Area Flux for Large Debris at Given Altitudes ................. 1.4 Debris at Given Altitudes and Inclinations ........................
4 5 6 7
Figure Figure
1.5 Evolution of a Satellite Breakup ............................................ 2.1 Active Netting System .................................................................
9 1 2
Figure Figure Figure
3.1 3.2 3.3
Conceptual Drawing of TV and Modules-Top View ....... 15 Conceptual Drawing of TV and Modules-Front View...1 6 Docking with Transfer Vehicle ............................................... 17
Figure Figure Figure
3.4 4.1 5.1
Conceptual Drawing of Netting Vehicle .............................. Mission Scenarios and Final Scenario .................................. System Breakdown .......................................................................
Figure Figure Figure Figure Figure Figure Figure Figure Figure Figure Figure Figure
5.2 Fuel Flow Configuration .............................................................. 5.3 Netting Vehicle Propulsion System ....................................... 5.4 Components of a Photovoltaic Space Power System .... 5.5 Communications Subsystem for the TV and NV .............. 5.6 V-Band Communications Network for the NV .................. 5.7 S-Band Communications Network for the NV ................... 5.8 Ku-Band Communications Network for the TV ................ 5.9 V-Band Communications Network for the TV .................. 5.10 S-Band Communications Network for the TV ................. 5.11 Tracking Range Characteristics ............................................ 5.12 GNC Integration System ................................................. .......... 6.1 Propulsion Module .......................................................................
Figure Figure Figure Figure Figure Figure Figure Figure
6.2 6.3 6.4 6.6 6.7 8.1 8.4 8.5
18 21 30 3 1 33 3 5 44 46 4 9 52 53 54 58 67 77
Netting Module (NM20/40/90 Configuration) ................ 7 8 Transfer Vehicle ........................................................................... 7 9 Debris Removal System ............................................................. 80 Launching Tube ............................................................................ 82 Docking Mechanism ..................................................................... 84 SPECS, Inc. Organization Structure ........................................ 90 Problem Solving with SPECS, Inc ........................................... 95 Manpower Estimates for SPECS, Inc ..................................... 96
viii
Table
of
Tables
Table 1.1 Orbital Debris Sizing Matrix ................................................. Table 5.1 Propulsion Requirements ..................................................... Table 5.2 Transfer Vehicle Electric Propulsion Options ............... Table 5.3 Electrostatic Thruster System ............................................. Table 5.4 Netting Vehicle Propulsion Requirements .................... Table 5.5 Primary Engine Characteristics .......................................... Table 5.6 TV Power Requirements ...................................................... Table 5.7 Battery Cell Comparison ....................................................... Table 5.8 NV Power Requirements ..................................................... Table 5.9 Thermal Subsystem Characteristics ................................. Table 5.10 NV Communications Characteristics .............................. Table 5.11 TV Communications Characteristics .............................. Table 5.12 Characteristics of the DPS for the NV and TV ........... Table 5.13 NV and TV Weight and Power for GNC ........................ Table 5.14 Summary of Volume Requirements (m3) ................. Table 5.15 Summary of Vehicle Masses ........................................... Table 9.1 Formulation of Projected Costs ......................................... Table 9.2 Anticipated Hardware Costs ................................................
ix
8 27 28 28 32 32 34 36 37 38 48 51 56 65 71 74 97 98
1.0
Project The
Most
Overview
problem
people
witnessed
have
Accordingly,
of
part
those
this
a difficult
of
detrimental
of
the
past,
SPECS,
Inc.
problem
orbital
effects
Inc.
launched
into
space,
requires these
objects
the
man
in
problem the
the
debris,
to
design
the
and
to
grasp.
much
space
philosophy
present
have
is
future
mission-related
less
environment. to create
an
problems
of
orbital
Voyager-2.
potential
The
fatal for
from
human
factors.
only
debris
began
[1.1,1].
in
1950's
this
first
because
current
to
payloads
in
place
lifetimes.
Since
the
escalated
Thus,
stages
and
recorded to
to major
only
when launch
technology
space,
orbital
booster
and
debris
because
consists
rockets,
In June
1983,
the
and
debris
.two
millimeters
from
other
example
shuttle in
Challenger
size.
The
can
do
the
space
shuttle
was
struck
estimated
of
anomalies
they of
be
TIROS-N,
These
A
a recent
reports
problems
contamination
however,
by
locate
TDRS-1,
mission;
illustrated space
ISEE-3,
vessels.
better
to
anomaly
range
pressure
degradation. is
examined A,
anomalies
spacecraft's
causes
be
ISEE-1,
punctured
the
can
In Appendix spacecraft:
debris
of
as
has
reports
debris.
shielding not
debris
space
upper
anomaly
on the following
thermal
resulting
orbital
fragments.
Spacecraft
found
defined
into
limited spent
has
orbital
rockets
payloads,
by
of
problem
payloads
caused
space
rocket
multi-staged
inactive
piece
pieces
of SPECS,
study,
inactive
Therefore,
orbital
is
debris. For
were
debris
seen
its
awareness
orbital
orbital
never
some
increased
of
indicate
damage incident. by
damage
a was
of
$50,000 to the space shuttle.
The untold damage was the danger to
the lives of the five crew members aboard the Challenger. Although
the dangers of orbital
debris had been expressed in
the past, this incident caught the attention of engineers and scientists in the United States and around the world.
Research projects with
the sole purpose of solving and understanding the problem of orbital debris
were created. Orbital debris is a problem because collisions
with tracked or
untracked objects can cause severe damage to both space vehicles and personnel.
Currently,
the chances of catastrophic collisions
small, but the chances are increasing. National Aeronautics
are
Dr. Donald J. Kessler of the
Space Administration
(NASA)
has voiced one
concern about the concentration of debris in Earth orbits.
He
theorizes that if the concentration of orbital debris becomes too high, there will
be self-perpetuating
collisions
among the debris.
called Cascade, or Kessler, Effect could result in millions untrackable
debris particles.
This so-
of small
[1.1,17].
SPECS, Inc. has initiated a program to clean up an important part of the orbital
debris problem.
Because Space Station Freedom
(SSF) is scheduled to go into orbit by the end of the decade, we have concentrated our attention on eliminating
debris near its orbit.
The
most feasible way of doing this is with an active system, a robotic spacecraft which collects the debris and removes it.
The following
sections presents the design process that was followed this goal. 1.1
Project
Objective
and
Scope
2
to accomplish
The project objective of SPECS, Inc. is to design a comprehensive
orbital
made debris falling centimeters. created.
debris removal
system that addresses man-
in the size range of one centimeter (cm) to fifty
To accomplish this objective, a project scope was
The project scope encompasses four major areas: debris
environment,
mission scenario, design options, and a debris
management
philosophy.
A mission scenario that efficiently targeted debris environment
addresses the problem in a
has been developed.
Feasible design
options that enable the mission scenario to meet its objective have also been determined.
Lastly,
a debris management philosophy
that
encompasses both the short term goals of SPECS, Inc., as well as the long term goals that have yet to be implemented.
1.2
Defining No
excluded the
last
area
of
from
the
30
areas
of
initial
tasks
of
Debris
the
years.
outer
contribute is
the
Environment
space
environment
barrage
of
objects
Although
space
are
the
more
SPECS,
Inc.
significantly
to
around that
debris
the
to
have
than
determine
debris
Earth
been
environment
populated
was
the
others.
huge, One
types
problem,
and
classified
as
been
launched is
what
has
some
of
of
during
the
debris
where
this
debris
located.
1.2.1
Types Orbital
debris
untrackable
debris.
Command
(NORAD)
of
Debris
may
be
The
North
presently
broadly
American tracks 3
about
either
Aerospace 7,100
trackable
or
Defense
objects.
Figure
1.2
displays the percentage breakdown
of tracked objects in space
[1.2,4].
Fragme_
o¢ is (46.s-/,)
It can the
Figure
1.1
Tracked
be
that
fragmentation
seen
tracked
that
an
present
objects.
additional in
1.2.2
low
50,000 earth
initial
around
this
the
provide it will
question
also is
debris
the
into
makes
tracked
- 60,000
orbit
Separated
up
objects,
objects
Groups
about
half
NORAD
too
small
to
of
all
estimates track
are
[1.1,13].
Location
environment
but
Besides
Debris A logical
Debris
SSF open
whether
choice the
Space
with up
for
Station
a protective
more
enough
targeting
options orbital
a debris
is
the
Freedom
(SSF).
Not
device
against
orbital
for debris
4
area
possible be
only
debris,
scenarios.
available
in
will
The the
vincinity
of the SSF.
Figure 1.2 displays a global outlook of all the
debris orbiting the earth [1.3,1.7]. that orbital
debris exists in most orbits around Earth; however, the
orbits containing naturally
low
of
debris In
was
the critical
Global
Figure
1.3
shows
orbit
(LEO)
(km) found addition
considered
800 in
this to
when
debris are
areas.
1.2
Earth
kilometers
the highest percentage of orbital
Figure
Further,
From this figure, it can be seen
Outlook
of
a large
[1.3,3.22]. km
the
debris
population
A targeted
was
chosen
Debris
Problem
for
altitude
because
of
altitudes
range the
high
in
from
200
percentage
area. the
altitude,
targeting
the
a region
5
inclination in
outer
of the space.
orbital In
debris
Figure
1.4,
500
1000
Altitude (kin)
Figure
which
1.3
displays
the
various
inclinations,
degrees
and
These the
target SSF's
an
65
SPECS,
dimensions small
of
Inc.
is for
debris
is
and
Large
Debris
debris
at
that
the
debris
certain
be
Given
up
400
28
[1.3,3.12].
match at
and
around
population
located
Altitudes
altitudes
inclinations
inclination
will
at
ideally
kilometers
with and
at
degrees.
defines
Debris
a piece
than
medium
seen
altitude
Orbital
greater
orbital
a large
which
28.5
for
be
have
for
Sizes
of
may
degrees
environment
1.2.3
and
it
of
Flux
amount
areas
inclination
dimension
Area
fifty debris
anything
of orbital
debris
centimeters. range
smaller
6
from than
as large
Similarly, one one
to
fifty
centimeter
if its the centimeters, [1.2,5].
By
lndinati_n
11, a
7o
3O
io 1 ()ffi
10 s
!0 4
10
_ULudo _m)
Figure
1.4
Debris
analyzing
the
dangers
the
space
environment,
size
was
debris
is responsible
elements
of
orbital
hours the
elements
The to threat
specifically
can for
approximity
a serious
Earth.
each
Altitudes
size
a logical
pieces
NORAD
create
that
Given
group
choice
and
of
for
Inclinations
orbital
debris
targeting
poses
a specific
to
debris
made.
Large
the
at
spacecraft
days)
to
of
collision. address
large to
maneuver
pieces
pieces
have
plenty
away
from
debris.
other
Earth.
positions
of large
debris
Further,
from
the
a spacecraft
would
large
tracked
cataloging
7,100 of
threat
be
capable of the designs
and
debris
are
Presently
[1.2,4].
known, of
time object, are
orbital
they
Since do
not
communicating (on
the
order
with of
thus,
eliminating
available
that
When considering
small debris, it was found that the
technology of structural shielding can be used to alleviate most potential
danger.
fortified
with
Present technology enables spacecraft to be
structural
shielding
that protects the spacecraft from
debris hits of less than one centimeter.
Since the number of small
sized debris in outer space is approximately shielding solution
four billion,
the idea of
against these small pieces seems to be the only sensible [1.2,4].
On the other hand, both tracking and shielding ineffective
against medium size debris.
17,500 pieces of medium-sized
NORAD
techniques are
has estimated that
debris exist, and because they cannot
be tracked or shielded against, they represent the most eminent danger to operational
spacecraft [1.2,4].
Using the sizing decision matrix shown in Table 1.1, the three sizes were compared in the categories of existing protection,
existing
designs, and existing debris quantity (0 = lowest concern, 5 = highest concern).
The medium-sized debris was recognized as the biggest
threat to the space environment.
Table
1.1
Orbital
EXISTING PROTECTION LARGE
Debris EXISTING OESIGNS
3
3
3
5
Sizing
Matrix DEBRIS
TOTALS
QUANTITY
7
(_[2E, JOcm)
MEDIUM (Lcm_I_-_l_m
SMALL (_]Z]_, lc:m}
2
3
13
4
10
1.2.4 After debris
in
focused the
Targeted
Debris
examining
the
the
on.
orbital
overall
engineers
debris
problem
cleaning
up
debris
ranging
from
300
debris
in
of
the
future
breakups
1.2.5
most Figure
km
of
the 1.5
debris shows
debris
full 1000
area
may
shift
in
the
the
of
that
by
debris of
satellite
targeted
more
was
progress
on
at
1.5
on altitudes
28.5 ° + 30 °. breakups.
Most
However,
area.
Breakup break-up area
is
a breakup
must the in
be
result several
assessed of
such stages
19].
Figure
sizes
area
concentrating
sized
from
and
a specific
at inclinations
targeted
history
felt
gained
Satellite
a satellite the
Inc.
of medium
resulted
of of
be
km
inclinations
environment,
could
this
dynamics
altitudes,
at SPECS,
clouds
Dynamics The
different
orbital
The
Environment
Evolution
of
9
a
Satellite
Breakup
since events. [1.4,15-
of
A satellite breakup debris cloud initially around the original differentials
forms an ellipsoid
location of the orbiting object.
Due to
of the particles in their orbital period this ellipsoid
evolves into an irregular, torus typically
narrow torus encircling
the Earth.
This
closes after several months to a year [1.5,223-241].
Further, the regression rates of the right ascension cause the torus to eventually
dismantle into a band about the Earth. This low density
band is limited altitude
in latitude only by the maximum inclination,
and in
by the extremes of the cloud. This phase is reached several
years after the event. The rate at which these phases are reached is largely a function of the velocity the initial
1.3
altitude and inclination
General In
project were
imparted to the debris fragments upon breakup and
Project
order
to
instituted.
following
and
amount
of
for
inclinations,
requirements Inc. the and
mission
mission: the
ability
are
listed
of
efficiently and
performance
design the to
meeting
team
ability capture
parameters
has
to
set
reach
a
the
the targeted
significant
debris
Performance rating
feasibility
SPECS,
requirements
altitudes
the
mission The
satellite.
Requirements
ensure
objective,
of the original
parameters
next
to them
•
Fuel
•
Power
•
Weight
( 0 = low
criticality,
Budget Requirements
10
below
with
5 = high
a criticality
criticality).
Criticality
5
Criticality
5
Criticality
5
1.4
•
Safety
Criticality
5
•
Resupply/Maintenance
Criticality
4
•
Lifetime
Criticality
4
•
Effects
Criticality
4
•
Cost
Criticality
3
•
Design Complexity
Criticality
3
•
Time
Criticality
1
on Environment
Constraints
Assumptions Along
with
mission requirements,
some general assumptions
were made to ensure a workable mission. •
Satellite
breakups will
eventually
•
Tracking
technology will
exhibit
torus qualities
accurately track orbit debris
down to a size of one centimeter (cm) •
With the help of cameras, the geometric shape of the debris will
be discernable
Other assumptions concerning will
have to addressed after further
2.0
Design
the
debris types
orbit. problem
research.
Approach
Initially,
all
the mission operations of the design
SPECS,
problem; of
debris
However, was
Inc.
considered
therefore, located
during narrowed
design from
the down
low
all
earth
medium
region.
11
of debris
scenarios
conceptual to
sizes
orbit
design
were
conceived
to
geosynchronous
phase,
debris
in assessing
within
the
scope a
for
of
targeted
the
2.1
Design
Options
SPECS, Inc. considered several designs to attack the problem of orbital
debris.
In considering
designs to capture debris, the tumbling
motion of the debris caused a problem when trying to grapple the debris directly.
However, by using nets to capture debris the
problem of spinning and tumbling debris is eliminated.
SPECS, Inc.
has designed an active netting system that uses Kevlar nets to capture pieces of medium-sized 2.2
Primary
debris.
Design
The active netting system shown in Figure 2.1 is composed of a Propulsion Module and a Netting Module.
The Propulsion Module is
used to perform the orbit transfers around the debris orbit and the netting
module performs proximity
maneuvers to reach the target
debris.
Each Netting Module contains several nets capable of
capturing debris sizes ranging from 1 cm to 50 cm.
•m,-a.t,-.g 'u_.,,.-
\
IIEL
/ Figure
2.1
Active
12
Netting
System
The active netting system will
target areas of satellite
where a high density of debris exists. netting system will 2.4
Design
breakups
Details about the active
be presented later in this document.
Philosophy
SPECS, Inc. has developed a debris management philosophy assult the present and future problems caused by orbital
to
debris. The
objectives of SPECS, Inc. have been divided into near term, mid-term and long-term •
strategies:
Near Term Strategy •
Attack the medium sized fragmentation
•
Develop an active system using a netting device (or similar
•
Reduce the collision
•
Single vehicle released from shuttle, space station or launched
•
probability
debris (1 cm to 50 cm)
in the target altitude
range
from Earth
Implement
an international
prevention
policy
on space
debris •
•
Mid-term
Strategy
•
Develop
a network
of active/passive
•
Launch an operational
•
Perform
orbiting
devices
base
area sweeps and explosion
clean-ups
Long term Objectives •
Increase operational
range to geosynchronous and transfer
orbits •
Expand system to remove the larger, tracked debris
13
design)
SPECS, Inc. realizes that correcting the problem of orbital debris is very costly and that the immediate satisfaction of cleaning up a few debris clouds will problem.
not have a noticeable effect on the overall
However, SPECS, Inc. has initiated a start toward solving
the orbital debris problem.
Hopefully,
other groups will join in
helping complete the SPECS, Inc. debris management philosophy, thus, solving the problem of orbital debris
and making the space
environment safe for the people of Earth and those wishing to visit.
3.0
System
3.1
Debris The
debris
Concept Removal
Debris
and
could
be
debris
Removal
capture
consideration,
it.
SPECS, used
System
to
System
Because Inc.
retrieve
as possible,
the
The
Transfer
Vehicle
near
a debris
torus;
orbit
Transfer
the
divided
two
components
Module
(PM).
Propulsion nets Netting
to
capture Modules
debris, docked
to
a net
was (TV)
carry
Each
Netting
the
Transfer later
14
shot
divided
into
and
the
many
then
Netting
Module
(NM)
Module
will
Vehicle use.
of
will
vehicle
pieces
two
of
main Vehicle
Vehicle
will
piece
a main
Netting
Netting
The
each
from as
Vehicle
debris.
out
under
capture
the
: a Netting
it for
seek
of debris
to
Netting
it seeks
and
size
that
vehicle
will
base
into
while
the
Vehicle
a temporary
actively
In order
main
the
of
decided it.
components:
will
to
use
a parking
the
Vehicle and contain have
(NV).
TV is
also
a several several
as
3.1.1
Transfer The
several near docked and
Transfer
Netting the
the
Vehicle
Modules
debris to
Vehicle
torus. front
will from
The of the
carry
the
main
Propulsion Transfer
the
Propulsion
base and
of
Module
operations
Netting
Vehicle
as
to
Modules
shown
in
and an
will
Figures
orbit be 3.1
3.2.
Net
Modules
Transler
Propulsion
Figure
3.1
Conceptual
Vehicle
Modulo
Drawing
15
of
TV
and
Modules-Top
View
Unspent Net Modu
_
Spent Net Module'.
Figure
3.2
After will
the
as
shown
the
departing
collect
all
spent
refurbished
Conceptual
debrisin
nets in
have Figure
Netting Netting
Drawing
from the
been
the
Transfer
prearranged expended
3.3. Module, Vehicle
The
of
Vehicle, atrget
area.
dock
with
and Propulsion
dock will
with leave
16
TV
Module
a new for
one, a new
and
the It the
Modules-Front
Netting will
Vehicle
return
Tansfer
when
Vehicle
will
separate
and
then
collection
View
from
the sweep.
Unspent
Net
Modules
=
Spent
Figure
Because several
3.3
the
months
Transfer
while
need
to
be
need
to
maintain
well
as
a consistent
3.1.2
able
fit
together
to
several
holes
separate
net
to
itself
a constant
for
Transfer
have
Vehicle
Modules
to
is
Vehicle
stay
in
collecting
orbit
for
debris,
it
will
an extended
period.
It will
also
especially
during
docking,
as
attitude,
orbit.
Vehicle shows
form in
will
Netting
to power
3.4
with
Vehicle
the
Netting Figure
Docking
Net
the
the
how Netting
Netting
capture
the
Propulsion Vehicle.
Module,
a single
piece
17
each of
and
Netting
As
can
of
which
debris.
be
seen, will
During
Modules there contain normal
will are a
operating conditions, these holes will being launched and retrieved,
be covered, but when a net is
the cover will
be retracted.
RCS (Reaction Control System) thrusters are shown on the Netting Vehicle in the figure.
These, or control moment gyros, will
needed to make adjustments to the Netting
Vehicle's
orientation
when the net is being retrieved so that it does not wrap around the spacecraft.
Netting Module
Net
/
Propulsion
Figure
3.4
Conceptual
Drawing
18
of
Netting
Module
Vehicle
be
Because the Netting Module is designed to capture several pieces of debris per mission, each mission will amount of fuel.
Considering that there will
require a substantial
be several Netting
Modules on the Transfer Vehicle, it would be very inefficient all the necessary fuel on the Propulsion Module. will
to store
Instead, the fuel
be stored on each Netting Module, and, when the Propulsion
Module docks with the Netting Module, a fuel link will
be established
between the tanks on the NM and the engine on the PM. Since the subsystems requiring on the Propulsion well.
the most power will
Module, the power system will
be located
be located there as
It will consist of a solar array mounted on the body of the
spacecraft and a rechargeable battery.
The power system will
linked to the Netting Module during docking in a similar
be
manner to
the fuel system so that the netting subsystem can be operated.
4.0
Mission Based
for
the
have
on
debris
been
Project
Scenario
mission
removal
evaluated
Mission
system
section.
are
designing
considered. seen
also
possible
been
considered.
have
the
options,
several
the briefly
criteria
listed
mission All
under
After
considering
final
scenario
scenarios
the
and was
scenarios
General
evaluating
chosen.
all
Altenative
discussed.
Options In
be
design,
using
mission options
4.1
can
netting
Requirements
reasonable
been
the
from
the
Alternatives the
logical
mission for
scenarios, different
structure 19
several stages
in Figure
4.1.
in
options the
have
scenarios
Arrows
in the
logical
connections between the mission elements indicate
loops.
The processes within a loop can be repeated until it is
necessary or intended to exit the loop. highlighted
closed
The connection between the
mission options indicates the final
mission scenario that
was chosen.
4.2
Final
Mission
4.2.1 The Modules bay
of
Active
Debris
Transfer
Vehicle,
will the
be
4.5
payload
weight
shuttle
for
altitude,
circular the
in the
debris
Remote
Removal the on
diameter Space
parking
Space
of
160 and
system orbit,
Manipulator
Shuttle
flights.
will
meters
System
be
being
three
is able
taking
placed
The is
the in
unloaded
(RMS)
payload
to carry
off in
Netting
The
long.
considered
after
20
and
when
been
Launch
Vehicle,
m 2 and 18
Shuttle
have
removal
System
Netting
two
an area
specifications
The
by
has
meters
process.
bay
launched
Shuttle
payloads
The
Scenario
of the
maximum
29,500
kg.
design a 400
from
km the
Shuttle.
cargo
I I
i
I
1 VEHICLE TV
I
TVINORBIT DEBRIS
I WITH DEBRIS I TV RENDEZVOUS
I I
FROM NET TV EJECT
I
mira CAPTURE
+
RETRIEVAL
I
I
I°_°._.°_iI_°_°N_ I i I
[i!!i!!!!!i .....................................
I ._
l',_i_,_ _iil _ F!_:!:!$_-::!::_.i:i_:_.::-.'-::; ::::::::::::::::::::::::::::::::::::
OFTV ORBIT CHANGE
" I
• I I IN STORE "JUNKDEBRIS ORBIT"
Figure
.................. _.........................'
,
iSTOR_ANAL_E _!!.i!'!iiii'_'_':: _' ....... _EO_IT_ODO'_SlIDEO_B_MO_ pY_RO_OLS'ON II BYTET_ER I 4.1
Mission
Scenarios
21
and
Final
Scenario
4.2.2
Orbital The
main
the
Propulsion
the
targeted
system
engines
The
Transfer
Vehicle
determined
from
lower
orbit.
Once its
so
Vehicle. panel
(50
100
the
attacted
aligned
one
Rendezvous The
Netting
Transfer
of debris.
when
the
size
orbit
the
piece
of 50
mechanism m.
The
to net
Transfer with
the the
a
actual
debris
parking
orbit
panels
will
along
the
side
of
Transfer
to
go
will
orbit be
of the an
the
the
Propulsion
and
capture
transfer and
used
to
one
the
debris.
track
move
the
net
reeled
the
will
a
debris
be
debris back
from with
According
when and
to
rendezvous
vehicle.
adequate
debris
be closed 22
orient
Capture
debris
km
the
be
it will
release
a Hohmann
targeted,
will
the
the
solar
use
catch
electric
the
position
an
been
orbit
in
and
All
by
position
Module
a few
of
its
center.
than
will
the
vicinity
altitude
has
axis
it
sensors
is within
of
to
carry
off,
Debris
will
Onboard
debris
a spring
a range
Vehicle
Vehicle
piece
detected
and
to
removal
using
a parking
this
there
Netting
km
performed
stabilized
From
1000
orbit
used
the
debris
control
into
reached
to
which
semi-major
gradient
sun.
to
go
is turned
are
a gravity
towards
4.2.3
they
be
will
has
system
that
also
the
be
the
its
orbit,
by
km)
vehicle
up
target
will
of
about
spiral
tracking
Vehicle
to
to
will
Modules
range
400
the
Tranfer
the
In
Module
by
ground
propulsion
retracted
reach
Vehicle
Netting
will
changes
The
slightly
the
to
inclination
Vehicle.
from
o _+ 30 o.
TV
Transfer
three
of 28.5
system
the
operational
orbits
propusion
major
and
The
comprises
of
of the
Module area.
inclination
and
Transfer
into
to
the
ejected is
within the
Netting Module by an attached chord. The Netting Vehicle will
then
target another piece of debris and go into a drift orbit to rendezvous with it.
This procedure can be repeated until all the nets of the
module have been used.
4.2.4
Netting
Vehicle
Resupply
The Netting Vehicle will resupply. Netting
It will
then return to the Transfer Vehicle
dock with the Transfer Vehicle to unload the spent
Module.
Another docking procedure will
provide
Propusion Module with a new Netting Module allowing Vehicle
to leave for another collection
4.2.5
Further After
the
same
the
Orbit
resupply orbit
in
Transfer
Vehicle
can
take
Modules
to
Netting
will
then
and to
remain Vehicle
target
Modules
are
return
Modules return
another
the
in
to
the same
new
order
or
to
the
filled Space
the
NettingVehicle
capture
further
procedure
with
debris.
Station Module
orbit
to
orbit wait
Resupply
for
Modules.
Base
on
SSF
and
or the
to
the
Netting
repeated
until
The
Transfer
Vehicle
the If
Propulsion
the
return procedure
return
is
the
This
23
all
attached.
propellant. 4.2.6
with
can
debris,
Module
This up
a similar
Netting
the Netting
sweep.
Propulsion
orbit.
Propulsion
a parking
and
the
the
Changes procedure
debris
the
for
spent the
TV
Netting is going
Module of
all
can
the
Transfer
will
save
to
For maintenance and resupply reasons, the active debris removal system will be based on the Space Station Freedom.
The
Space Station is assumed operational by the year 2000. For these operations, the Transfer Vehicle will certain area within the proximity
first fly into a
of the Space Station.
Using EVA
astronauts or the robotics on the Space Station ("Canadarm" servicing will
system), any maintenance that the Transfer
be performed.
three new Netting Shuttle.
The Transfer Vehicle will
mobile
Vehicle
needs
be resupplied with
Modules, which have been launched via the Space
Spent Netting Modules will
bay for return to earth.
be placed in the shuttle payload
If the shuttle is not available, the spent
Netting Modules will be attached to the Space Station truss at a predefined
area, where they are stored until they can be deployed
to
the shuttle payload bay. This resupply option seems to be reliable: the shuttle will supply the Space Station frequently, Modules will Station.
and space for the Netting
be available with the logistics modules for the Space
For the reentry purposes, the maximum
payload reentry
weight for the shuttle (23,000 kg) must be considered.
4.3
Discussion Because
sequences, less on
maneuver system
of
Alternative
difficullty
a free-flying
reliable this
of
than
base of proves
the to
base
the
would
Space
be Shuttle be
in
Missions the
was
and
considered
to
Station
necessary, would
effective,
resupply
and be
24
be
more
based
system:
additional
an
additional
docking
required.
SPECS,
maintenance
Inc.
Nevertheless, considers
complex
and
robotics
if
expanding
the the
active debris removal flying
resupply
system to different
base with
inclinations
using a free-
onboard robotics.
Using only one vehicle for the orbit transfers and the debris capture operations was not considered to be efficient
because of the
large amount of fuel needed. The option of transferring
the Transfer Vehicle
debris orbit was excluded due to safety aspects. probability
with debris is relatively
directly
into the
The collision
high and poses a high risk,
especially for the large solar panels. Deorbiting
the spent Netting Modules from the debris orbit by
either a disposable Propulsion
Module
or an additional
has been excluded due to cost and mass.
deorbit
device
It was decided that a
reusable system would be cheaper in the long run, and would also limit
the production
of further
debris.
For comparable reasons, deorbiting
the Netting Modules
by
special deorbit devices from the Space Station was considered less effective
than the deorbiting
this scenario will
with the Shuttle.
On the other hand,
be strongly dependent on the ability
Shuttle flights
to send used Netting Modules to Earth in the Space Shuttle. Therefore, the storage of the spent modules in a safe orbit that could be tracked from Earth is still a viable option. The deorbit of the modules by a tethered deployment
and
release has also been considered since this system is being designed for the Space Station to use with reentry capsules. If the tether system is operational removal
by the year considered to launch the debris
system, this option
can be reconsidered.
25
The final option of storing and analyzing the debris on the Space Station seems to be feasible, but safety aspects as well as the minimization
of Extra Vehicular Activities
(EVA)
have to be
considered.
5.0
Subsystem The
General
list
mission
requirements
Requirements
that
section
for
all
subsystem
•
Propulsion
•
Power
•
Thermal
•
Communications
of
selections.
were
this
defined
report
These
in
the
formed
the
subsystems
include:
Control
•
Data
Processing
•
Tracking
•
Guidance,
•
Netting
•
Structural
•
Fuel
System
Navigation,
(DPS)
& Control
(GNC)
Materials
Requirements
Propulsion
5.1.1
Transfer Electric
transportation electric
of
Project
foundation
5.1
Design
propulsion for
propulsion
requirements
Vehicle
as
the was shown
was
selected
Transfer made in
as
Vehicle. under
Table
5.1. 26
a list In
the
method
The
decision
to
of
specific
propulsion
comparison
of
with
select
chemical
propulsion,
electric provides greater efficiency,
lower fuel costs,
greater operating times, and a lower chamber pressure for easier fuel storage [5.7].
NASA prohibition of H202 in the Shuttle Bay limits
chemical propulsion
choice to monopropellants
[5.5].
The low-thrust
option is a viable choice for the Debris Removal System since time is not a critical
factor.
Table
5.1
TableS.
Propulsion
Requirements
1 Propulsion
• Clean
Exhaust
• Storable eFuel
Requirements
Fuel
Production
eTotal
Cost
Propellant
Mass
• Time Of Flight • Thruster Efficiency eOperating
Time
eFeasilbility
Selection through is
a compiled of
ISP,
selected of
list
the
With
total
the
a comparison
each
high
of
of
specific
type
of
different
several
four
thruster
a low
input
an
electro-static
as
the
primary
10
ion
thrusters
four
of
electric
thruster
propulsion
important
was
made
systems.
performance
Table
characteristics
5.2 on
classes. power,
good
Xenon
ion
propulsion will
be
thruster propulsion
unit used
27
efficiency,
for to
the
provide
system Transfer the
and
a fairly
has
been
Vehicle. continuous
A
thrusting
that the Transfer Vehicle
needs to reach the target orbit,
which is about 50 km. below the debris torus.
The ten engines will
Table
5.2
Transfer
Vehicle
Electric
Propulsion
Options
Table
5.2
Transfer
Vehicle
Electric
Propulsion
Options
Thruster
Thrust
Power
Life (hrs.]
(kw)
ISP
Arcject (NH3) Ion (Xe)
968 3600
37% 70%
2.4
Ion
7000
80%
5.4
50.0
5000
30%
I00.0
500.0
MPD
(s)
(N)
Types
Efficiency
(A) Thruster
continue the
regenerative
breakdown the
0.4
[Reference
5.7,9,10,
even
shadow
thrusting
during
MnO2-H2
of
Transfer
the
30.0
batteries.
electro-static
3-5
5.3
Table System
Table
Electrostatic 5-3
5.3
through is
a mass
propulsion
Mass:
• Total • Total
Power: Volume:
power
from
and
power
onboard
Electrostatic
2550
System
Thruster Total
System
Mass
not
35 kw. 5 m'3*
10 only
include
ion 4 N
of
or power
onboard
Transfer
thrust.
I000 120
kg. kg.
630
kg.
800
kg.
• Radiator
fuel storage
thrusters
Breakdown:
•Thrusters ePPU
kg.
sCradle
generate
1,000
system
Thruster
Configuration:
oTotal
The
lO, OOO
Vehicle.
Table
*Does
10,000
l 1,13l
times
Xenon
750
The
28
the thrust
Mass supply
Vehicle
acceleration
of
volumes.
will the
Transfer
Vehicle kg.
will
be about .0002 m/s2 for a system mass of about 21,000
With such a low acceleration, concern arises as to whether the
thrust acceleration solar to
pressure.
can overcome perturbations
Appendix
determine
the
C contains
Transfer
such
a number
Vehicle's
of
acceleration
as
drag,
J2,
or
computations
due
to
done
these
perturbations. At 4,300
a station
times
smaller
The 5,000
computed
times
smaller
The Transfer the
largest
for
and
periapsis
eccentricity
operational degrees
heat
the
induction
the
thrust
is
low
kg/m 3 for into
only
at a pressure to
be
density
Xenon
gaseous
vehicle
was
about
8 times from
so
the
that
smaller is
J2
only
misleading
eccentricity
change
a periodic
than
in
change
the in
nodal
the
[5.1,14].
The
needs
about
by
a secular
diagram
storage
liquid
was
directed
but axis
the
experienced
applies
[5.11]. and
on
perturbation
J2
was
acceleration.
large
a functional
Xenon
a very
acceleration
which
J2
elements,
Celsius
Xenon
to
perturbation
acceleration.
thrust
change.
is
drag
perturbation
semi-major
5.1
the
thrust
The
orbital
The
has
3520
than
configuration
atmosphere.
to
pressure
transfer,
and
Figure
Xenon
solar
due
axis
km,
the
was
orbital
semi-major
400
than
acceleration.
and
111
of
computed
Vehicle
thrust
since
orbit
of
Xenon
stored
chamber.
29
less
thruster as than
as a liquid
of about
before
ion
is stored
slightly
[5.3].
form
the
5 kg/m 3
A vaporizer it reaches
a liquid one since as
gaseous compared
is employed the
at -
electric
to
Mass
Flow
0= .4
P
= 250
Torrs
T
= 300
Degrees
Xe+
-->
35
km/sec
C.
kg/hr
Power
Pump
(35
Figure
P
=
1500
Torr
T
=
-111
Degrees
5.1
System
30
Input kw)
C.
Breakdown
II
Main Fuel
Line
Pump
! I I l I l
I
'
Distribution
Distribution
Node
.^..
sssoss
I
Node
I••SOS
I
i
Figure
i
I
5.2
!
Fuel
31
Flow
Configuration
_
I
I
I
I
5.1.2
Netting Selection
different
set
Table
Space 5.5
5.4
• Low
Power
Netting
Vehicle
e Storable
Propellant
bi-propellant,
a list
of
the
in
Table
required
5.4.
Propulsion
Requirements
Propulsion
Requirements
with
highest
Hydrazine-N204,
has
Module been
engine/
has
[5.2,5].
chosen
been
The
for
propellant
Isp
the
possible
selected
Primary
Netting
to
Vehicle.
characteristics
Table
5.5
Primary
Engine
Table
5-5
Primary
Engine
Propellant: Chamber Pressure: Oxidizer/Fuel Ratio: Thrust (Vacuum): Restart
Characteristics HydrazinelN204 7.5 Atms. 1.6:1 N
3 18 sec.
Lifetime:
• 20,000
Gimbal:
times
6 dgs. / pitch - yaw 1.168 m.
Width:
1.958
Length:
32
of the
in
Characteristics
3870
the
Table
utilized
(Vacuum):
fuel
Engine
Vehicle.
ISP
a
Capability
Propulsion RCS
shown
system
Consumption
Engine
the
propulsion
Availability
e Throttle
on
as
Vehicle
Netting Thrust
Vehicle
criteria,
Netting
5-4
Propulsion
Netting
propulsion
Shuttle's is
the
eHigh
A engine
of of
Table
Vehicle
m.
the
Fuel/Oxidizer Storage Tanks GNC Control Lines
Pumps
Throttle
Figure
5.3
Netting
Station
Vehicle
33
Propulsion
System
Figure Vehicle's as
the
5.2
5.3
primary
functional
system
diagram
with
of
the
connections
to
Netting GNC
as
well
mechanisms.
Power
Transfer
The
be
the
propulsion
gimbaling
5.2.1
the
illustrates
main
propulsion used
power
by
the
moving
parts,
this
power.
The
power
are
Much
Transfer
are
Power
consumer
subsystem.
photovoltaics
Vehicle
Vehicle
of
the
Transfer
of
the
power
that
ion
thrusters.
Vehicle's
electric
a clean
source
of
power
source
was
requirements
shown
in
Table
of
Table
5.6.
5.6
TV
Supply
energy
with
chosen each
Power
Vehicle
no
to
is
system
is
produced Since
will solar
mechanical
supply
subsystem
the on
the
needed Transfer
Requirements
Communications
280
DPS
50 W 260W
GNC
W
35kw
Propulsion Structures
13W
Thermal Tracking Total
Because the
Earth's
the shadow,
solar
photovoltaic
batteries
will
power be
34
needed
35.6 kW
system to
can
power
not the
operate vehicle
in
during this time.
For one cycle, a 1.5 hr. orbit , the shadow time is 36
min. [5.14]; therefore, the battery charge time is about 58 min.
To
power each system during each cycle, various sections of a power system are needed. estimates
of
As
several
shown
power
in
Figure
components
5.4,
mass
are
required.
Power Management and Distribution
and
power
_
To Load
Storago System
Figure
5.4
Components
of
sunlight
times,
During subsystem Using
and
charge
a specific
power
approximately
900
and
weights
PMAD have
structural system. a reduced
the
kg.
a
Photovoltaic
the
arrays
batteries,
of
100
will
which
W/kg that
this
[5.17]
and
estimates
array
sensitivity
to
technology radiation.
35
need is
[5.16],
Note
Current
Space
to
value
of
arrays
includes of
the
includes 1997
System
power
a total the
By
Power
each
about
73
will
weigh
both
blanket
radiator solar cells
and
kW.
the
cells
that
should
be
[5.15]
available that are completely
tolerant to radiation
[5.18].
The array
size is needed to better define the appearance of the Transfer Vehicle and determine the mass.
For the required system power, the
two arrays must have an area of 76.2 m2 with dimensions
of about
5.0 m x 15.24 m, with one array on each side of the Transfer Vehicle. This array size is calculated assuming that 35% efficient cells will
GaAs solar
be available by the time that the vehicle is in operation.
Several batteries were acceptable for the Transfer power requirements,
Vehicle's
but the one with the best combination
of
characteristics is a new development for high-cycle life LEO, rechargeable
MnO2-H2 cells [5.19].
energy, used in determining
This battery has a high specific
the mass of the battery, as shown in
Table 5.7 ompared to the widely used NiH2 battery.
Table
5.7
Battery
Property
Cell Ni-H 2
Comparison Mn02-I'I2
Specific Energy 22.2
(Wh/kg) (Wh/l)
33.1 82%
78%
8O%
85%
10,000
25,ooo
Efficiency Maximum
DOD*
Cycle Life *Depth
For discharge
LEO of
the
of Discharge
applications battery-is
the
cycle
necessarily
36
life-one
orbit
with
a charge
large.
The
efficiency
of
and this
battery is comparable to the widely used Ni-H2 battery. interest is the large depth-of-discharge,
Also of
85%, which is the maximum
amount of energy available to be drawn from the battery per cycle. Using the specific energy of the MnO2-H2 battery, a weight of about 741 kg is found with a volume of 0.721 m3.
5.2.2
Netting The
the
subsystems
same
Transfer much
Vehicle
type
of
of
the
Supply
Netting
Vehicle
photovoltaic
array
The
power
requirements
the
Transfer
Vehicle
Vehicle. smaller
Power
than
Table
5.8
NV
and
will
battery of as
Power
Propulsion Structures
12W
Thermal
13W
2.4
calculations and
battery
weighs
is
m 2 which
propulsion arrays
module will
be
run
less
of
made
each 13 kg
calculated is small
Table
are
5.8.
wm_
were
battery
weight
in
Vehicle
260W
Total
array
Netting
the
5ow
Tracking
MnO2-H2
for
107 W
GNC
the
the
with
Requirements
DPS
charge
powered
selected
shown
Communications
Similar
be
as
the
exposed
and
has The
to mount
Netting to
find
subsystem
14 kg.
enough
to
561
debris
37
the during
a volume required the
Vehicle.
W
solar In this
impacts.
necessary sun
power
times.
The
of 0.01265 area cells way
m3.
of the
The
array
directly the
to
solar
on
is the
5.3
Thermal
5.3.1
Subsystem
Transfer The
passive
Thermal
cooling
paint
and
heat
dissipation
plates.
dissipated
through
are
mounted
cooling to
used
on
Freon
the
circulation.
5.9
Thermal
Table
(kg)
Power (W) Volume (m^3)
5.3.2
Netting The
NV
Module
(NM)
thermal
control
outer The
thermal network system
structure outer
Thermal
hull
and
Table
active
consists
of
is
TV
Subs,_stem
as
heat
the
heat
the
is
controlled
through and
and
radiation
such
excess
5.9
the
via
an
inner
exchangers
contains
the
thermal
weight,
subsystem.
Characteristics
I_M
PM
TV
436.0
235.0
303.0
7.0
5.0 .85
13.0
3.95
2.75
Vehicle subsystem
can
be
and
the
Propulsion
for
the
NM
maintaining
is protected
dissipation
Pumps
for
Characteristic Weight
the
circulates
perform
requirements
and
environment
temperature.
both
equipment,
inner
the
power
requires
system
plates
maintain
and
TV
External
The which
the passive
these
them.
system,
to
volume,
for The
are
volume
subsystem
networks.
antennas,
active
Vehicle
plates
into
Module
solar
passively
38
fuel
for
off
by the
Netting network.
cooling
temperature
radiation bleed
the
(PM)
is responsible the
from
divided
the
requirements. special
excess
paint. heat
The
generated by the externally tracking devices).
attached instruments
(OMNIs
The internal NM environment
and
requires an active
cooling system to maintain the temperature of the stored fuel.
This
thermal network uses a Freon cooling loop with a pump to circulate the fluid
and heat exchangers to regulate the temperature.
thermal network dissipation
also uses the radiation
plates to passively
The inner environment
control
paint and thermal
control the outer hull
temperature.
is regulated by an active system similar
the system used on the NM. temperature,
The PM
In addition to controlling
the fuel
the PM network controls the guidance, navigation
subsystem, the data processing subsystem,
communication
computer temperatures.
to
and
and the
However,
the PM network
is
not as extensive as the NM network because less fuel is stored on the PM as compared to the NM. and power characteristics Module
5.4
thermal
of the Netting Module
and Propulsion
subsystems.
Communications
5.4.1
Subsystem The
certain
require
design
of that
provide
needs
of
Requirements the
vehicle
on
the
requirements
separation
to
Table 5.9 contains the weight, volume
to be
communications
the a
Netting
of to
Vehicle
link
(NV) link
data.
control
the
operational
communications
communications
transfer able
and
In the
developed
be
addition, NV for 39
via the
sequence
subsystem.
and
the
the another debris
The
Transfer
established
Vehicle
between
remote link. removal
levy
the
command
(TV) two center
The system
must
be compatible will
with the STS Orbiters because the resupply sequence
be conducted with these vehicles.
During resupply it may be
necessary for command of the TV to be handed over to the STS. Thus, the communications
subsystem on-board the TV must be
capable of communicating
with the STS Payload Interrogator
Also,
the
TV
Nominal
must
operations
TDRSS,
to relay
located
on
the
contingency
ground
orbital
positions could
or
with
the
types
requirements
be
this
endpoints,
data
have
command
Freedom
center
a direct
center
or
an
enough
Based
on
appropriate
in
each
time
these
the
by
the
general following
communications
subsystem.
Transfer
Vehicle •
•
The
TV
will
receive
external
control
The
will
TV
data
via
TDRSS
from
the
center.
transmit
external
control
•
The
TV
will
transmit
°
The
TV
will
receive
the
command
data
and
video
via
TDRSS
to the
center. command data
and
NV.
40
data video
direct from
to the
the
a direct
three
case,
for
In
be limited
given
developed
will
(SSF).
of
will
center.
satellite,
capable
required
been
the
command
capability
established. of
This Station
command
two
and
Space
the
of
be
center.
must
Although
with
telecommunications
TV
SSF). the
the
to the
the
eventually
requirements
communicate
utilized
or in
link
(STS
to
signal
operations
vehicle
specific
able
will the
communications
link
be
(PI).
the direct
NV. from
be
Netting
Vehicle •
The
NV
will
receive
•
The
NV
will
transmit
These needed
requirements
by
the
TV
developed
based
the
Station
Space
Vehicle
the
the
communications and
and
sample TV
the
the
to
guideline
TV.
outline
a subsystem
for
the
TV.
the
subsystem
designed and
from
direct
basic
these
(SSF),
SSF.
and
These
NV
the
Orbital
was
Orbiters
(STS),
Maneuvering
TDRSS
satellite
analog
S-Band
(1.7
to 2.3 GHz),
[5.20,559]. and
The in
a phase system.
of
data,
the
orbiter
to the
system
can
to
examine
the
OMV,
compatibility
that
is capable and
of
aided
receiving
video
over
(4 to 6 GHz),
this
satellite
of
operational
is limited
in
the
guidelines the
communicates
in
design
of
frequency
and
Ku-Band
range,
the
the
S-Band,
The
and
S-band
PM
system
is
ground
via
TDRSS
or ground
S-Band
a Frequency
used
to
Ku-Band,
system
from The
the
to 14
between
Modulation
stations. to
(12
bands.
communicate
directly
bands:
link
either
system
41
three
three
(UHF).
information
transmitting
these
(PM)
transmit
and
to
band
only
necessary satellites,
provided
Frequency
modulation
Orbiter
was
TDRSS
subsystems
C- Band
TDRSS
Ultrahigh
The
the
programs
audio
Because
STS
it
subsystems.
and
TV
criteria
subsystems
digital
GHz)
design
communications
The
and
video
the
subsystems
direct
Approach
developing
STS
the
and
From
Freedom
Design In
the
NV.
data
(OMV).
5.4.2
the
data
provide
and on
command
S-band
ground
includes (FM) the FM stations
in contingency operations or during Department of Defense (DOD) missions. with
The S-Band system is the means the Orbiter communicates
detached payloads.
The primary
the Orbiter is the Ku-Band receives information
system.
communications
system for
This system transmits and
to the ground via the TDRSS satellite.
Because
the TDRSS has a problem locking on to the narrow beam of the KuBand signal, the S-Band is used to establish antenna lock with the TDRSS and then the link is handed over to the Ku-Band system.
The
UHF system is the means the Orbiter communicates with the EVA astronauts.
This system is a voice link only [5.20,573-598].
Based on this system definition
it is clear the Orbiter would
communicate directly with the TV using the S-Band FM link. Therefore, to support contingency operations, the TV must be able to receive an S-Band signal from the Orbiter. capability
Additionally,
the
must exist to command the NV directly from either STS or
SSF, should communications between the TV and the NV be lost. Thus, the Orbiter would communicate to the NV using the S-Band FM link, and the NV must be able to receive, the transmission. communications
Any
with the TV via TDRSS would employ the Ku-Band
system. The SSF communications communications communication TDRSS.
system.
system is similar
Ku-Band is the primary
to the Orbiter means of
between SSF and the control center through
the
The Ku-Band system is also capable of direct communication
between the SSF and a vehicle with the line of sight. communications the UHF system.
within
a proximity
of 1 km will
Any direct
be completed using
SSF has an S-Band capability that could be used for 42
direct link in contingency operations.
Because of the similarities
between the Orbiter and the SSF communications and NV capabilities
systems, the TV
outline in the STS section remain unchanged.
As a possible sample design for the TV and NV, the OMV communications
systems was examined.
The OMV will
communicate
to the TDRSS satellites,the SSF, the Orbiter, the Deep Space Network (DSN) and the Ground Spaceflight Tracking and Data Network (GSTDN)
via and S-Band RF link [5.21,21].
requirements influenced the OMV,
These compatibility
the requirements for the TV system.
Like
the TV will be capable of communicating with the Orbiter,
TDRSS, and SSF. Because the TV will not be travelling out of Earth orbit there is no need to communicate with DSN.
Additionally,
because GSTDN is being phased out by NASA in favor of the TDRSS constellations,
5.4.3
this requirement
Subsystem From
an
requirements, 5.5
was
system
employed
to
capability
is
the
SSF.
The
56
GHz)
system overused
as
its
also TV
prevent band.
these
designs
V-
communicate
method.
the
Ku-Band
with
the
Band
TV
choice
was
to
in
the
design
shown
S-Band
TDRSS
The
over
NV an
signals
S-Band due
against
to
to the
in
will serve
Figure using
a
be as
a
S-Band with
via
system
TDRSS link
and
fail.
the
weighed 43
resulting
communicate
with chosen
was
An the
the
through
system
communicate
interference This
will
lock
on
and
subsystem
signal
needed will
TV primary
should
system. to
The
establish
system
of
communications
developed.
Ku-Band
backup
Design
evaluation the
was also unnecessary.
the
a V-Band or the
STS (46
and to
Ku-Band crowded
addition
of
and
another antenna and found valid.
Again the S-Band will
be used to
acquire signal lock and then the V-Band will take over. system will
then become the backup communications
The S-Band link
between
the TV and the NV. TDRS K-Bar_l
==IF=
S-Band
K-Band
9
S-Band
J_V-Band
STS
TV
SSF
@ Figure
5.5
The
Communications
TV
and
from
the
external
and
video
from
the
TV.
The
the
NV
will
command its
TV
for
required
to receive
be
center.
internal will
Subsystem
The
computers
then
transmit
NV and
the
the
will
and
and
command
transmit
external
data
TV
video
uplink
both
cameras to
NV
data
back the
to
control
center.
5.4.4
Netting The
Netting
communications means
of
(TV).
The
The
NV
Vehicle Vehicle
capability.
communication S-Band
network
communications
(NV)
will
The
both
V-Band
network
the
and
between will
have
be
subsystem
44
NV
V-Band will the
and be
Transfer
S-Band the
primary
Vehicle
used
in
contingency
operations.
will
be
responsible
for
transmitting
both data and video to the TV and receiving
command
data from the TV.
5.4.4.1
V-Band
Network
A schematic of the V-Band network is shown in Figure 5.6. low power, low gain, hemispherical (OMNIs) provide
will
receive and transmit the NV signals.
sufficient
and the TV. two OMNI
gain for proximity
antennas These antennas
zone operations between the NV
The OMNIs also have a wide TX/RX range;
thus, with
antennas mounted on opposite sides of the NV,
communications Additionally,
omnidirectional
Two
will
be virtually
independent of attitude.
OMNIs have a smaller surface area than the parabolic
antennas, which is desirable for NV operations in high density debris zones. The remaining
components of the V-Band network
Switch, the Transmitter-Receiver, Band Switch is an electrically the two OMNIs (VSP).
VT-R
driven
and demodulation
the power amplifier
signal, and the filters. network is the VSP. Distribution
The V-
switch that alternates between
when commanded by the V-Band Signal Processor (VT-R)
signals.
The
that regulate the carrier
that steps-up or steps-down
the
The "brain" of the V-Band communications This unit is receives input data from the Video
Subsystem (VDS)
and the Data Management Subsystem
The data is first encoded and multiplexed,
VT-R for transmission.
performs the
of the inbound/outbound
contains the crystal oscillators
frequency,
(DMS).
and the Signal Processor.
The V-Band Transmitter-Receiver
modulation
are the
Additionally, 45
then sent to the
the VSP receives signal data
OMNI
1
OMNI
2
Switch
V-Band T-R 1 • Oscillator • PA, f'dters • Modulates/ Demodulates
V-Band Signal Processor 1 • Encoded Decoder • Multiplexer/ Demultiplexer • Embedded Controller • Switching • Monitoring
t
_
Data (DMS) Video (VDS)
V-Band Communications
Figure
from
the
5.6
VT-R,
the
DMS
for
and
the
Switch.
V-Band
software
upon
is also
performed
Communications
demultiplexes
processing. All receipt by
The switching of the
Network
the VSP.
Network
and
decodes
VSP
also
Upon
46
data,
controls
commands command
the
are from
detection
for
then, monitors
initiated
by
DMS. of a fault,
NV
relays
and
the
the
the the
Fault the
it
VT-R
VSP detection VSP
to
software notifies the DMS. volume
characteristics
Table 5.10 shows the power, weight, and
of the V-Band
communications
The V-Band network is the primary between the NV and the TV. failure of one of the units will
network.
means of communication
The network is single fault tolerant: a disable the entire string.
of the V-Band network the NV can utilize
Upon failure
the S-Band network for
communication.
5.4.4.2
S-Band
Network
The S-band network has receive only capability contingency operations.
for
A diagram of the S-Band communications
network is shown in Figure 5.7. Again, an OMNI
antenna was chosen to provide maximum
coverage.
Only one S-Band OMNI will be located on the NV, thus,
inhibiting
communications
to certain attitudes.
Incoming
data is sent to the S-Band Receiver for demodulation.
command
The data is
then sent to the S-Band Signal Processor (SSP) for demultiplexing decoding.
and
The resulting command data is shipped to the DMS for
processing. The S-Band network is the secondary means of communication for the NV.
For contingency operations only command data can be
received by the NV.
Most likely, this data will direct the NV to
return to the TV from repair. significantly
For this reason the S-Band network is
scaled down when compared to the V-Band
47
network.
Table
5.10
ORU
V-Switch
NV
Communications Power
(W)
Subsystem Weight
2 (ss)
(kg)
Characteristics Volume
(m^3)
.907
.0018
15.87
.0183
8.16
.0117
.0018
25 (Switching)
V T-R
VSP
50
30
S Receiver
35
2.27
SSP
30
15.87
OMNI
N/A
20.41
.0117
.0006
107 (V-Band) TOTALS
140 65 (S-Band)
48
.0457
OMNI
• Demodulates ° PA, filters S-Band Receiver
1 S -B and Signal Processor • Deformatts
Data (DMS)
S-Band Communications Figure
5.7
5.4.5
Transfer The
on-board:
Ku-Band,
V-Band,
and
via
the
TDRSS
NV.
the
primary The
Control
(TV)
means
Center
of
network
has
and
communications
S-Band
Network
for
the
NV
Vehicle Vehicle
TV
the
Communications
Transfer
the
as
S-Band
Network
three
S-Band.
(CC)
communications
49
the
networks
Communications
is accomplished
satellite.
provides
communications
The
by
V-Band
between secondary
the link
between a Ku-Band
network TV
and
between
link
is
used
the the
TV and the NV and serves as the backup network for the Ku-Band network. Additionally, Ku-Band network. TV communications
5.4.5.1
the S-Band is used to acquire TDRSS for the Estimates of the power, weight and volume of the subsystem are provided in Table 5.11.
Ku-Band
The Ku-Band
Network network provides the communications
between the TV and the CC via TDRSS. network is shown in Figure 5.8.
link
A diagram of the Ku-Band
Because of the large distances the
signal must travel, 3 foot diameter, high gain, parabolic
antennas
were chosen for the Ku-Band network.
These antennas are
directional
through a two axis
gimballing
and have pointing
capability
mechanism.
The antennas are controlled by an Antenna Controller which regulates the motion of the antennas. to each of the antennas by an electrically gimballing
The ACON is connected
driven switch.
All
commands are issued via the ACON upon request from the
Ku-Band Signal Processor (KSP). for failures
(ACON),
The ACON also monitors the gimbals
and performs a small degree of fault detection on the
switch, gimbal motors, and itself. The remaining (KT-R)
components,
and the KSP, are functionally
the Ku-Band
Transmitter-Receiver
identical to the ST-R and the SSP
discussed for the NV. However, these units will designed for the Ku-Band frequency range.
be specifically
The KSP receives inputs
from and outputs data to the TV DMS and the V-Band network.
50
Table
5.11
TV
ORU
Communications Power
Subsystem
(W)
Weight
(kg)
Characteristics Volume
(m^3)
Switch
2 (SS) 25 (Switching)
.907
.0036
V T-R
120
15.87
.0184
VSP
30
8.16
.0117
ST-R
35
2.27
.0018
SSP
30
15.87
.0117
OMNI
N/A
20.4
.0006
KANT
15
27.2
1.601
ACON
30
15.87
.0082
120
15.87
.0184
30
15.87
.0117
KT-R
KSP
107 (V-Band) 65 (S-Band)
TOTALS
1.687
154.2
220 (gu-Band)
5.4.5.2
V-Band The
similar
from
V-Band
to
network
the
is the
Network
NV
communications
network
shown
in
components
used
on
Figure
network the
5.9.
NV.
is
the
The VSP.
51
A only Because
on-board
the
schematic
of
component the
TV the
that TV
operates
is V-Band differs as
a
KANT1
KANT2
Switch
ACON
I
1
KT-R
• Motion
I
• PA, filters
• Monitoring
• Modulation
L Ku-Band
SP 1
• Encoding Decoding • Multiplexing/ Dernultiplexing • Embedded Controller • Switching • Monitoring
Data (DMS) V-Band System
Ku-Band 5.8
Figure
relay
station
and
demultiplex
between
simply
shipped
and
to the
on
Ku-Band
the
the
CC.
Network
Communications
NV
incoming
through
Communications
the
Similarly,
and
the
data
CC,
stream
V-Band outbound
52
Network
there
is
from
the
network
to
command
no
for
need
NV.
the data
the
to decode The
Ku-Band from
TV
data
is
network the
Ku-
OMNI
1
OMNI 2
Switch
V-Band T-R 1 • Oscillator • PA, filters • Modulates/ Demodulates
V-Band Signal Processor 1 • Embedded Controller • Switching • Monitoring
System Ku-Band DMS
V-Band Communications Figure
Band
5.9
network
multiplexing and
monitoring
V-Band
is
Communications
transmitted
required.
Network
through
Essentially,
the the
functions.
53
Network
VSP VSP
with only
for
no
the
TV
encoding
or
provides
switching
OMNI
\7
S-B and T-R • Modulates Demodulates • PA, filters
S-Band Signal Processor • Multiplexes/ Demultiplexes • Encodes/ Decodes
Data (DMS)
2
S-Band Communications Figure
5.10
5.4.5.3
S-Band The
Band
network
receive
and
perform
network
used
network
for
transmit the
Communications
Network
for
the
TV
Network
S-Band
S-Band
to
S-Band
Network
on
for
the
the
TV.
the
NV.
.S acquisition
extensive
5.10
shows
TV
S-Band
has
The of
is more
Figure The
signals.
TV
transmit
signal
54
function
the
a schematic the
capability for
than
the
of the
capability was
S-
to necessary
Ku-Band
network.
After acquisition
over to the Ku-Band
of signal, the S-Band network will
hand
network for data transfer.
The components of the S-Band network consist of a single high power, low gain hemispherical OMNI Transmitter-Receiver
antenna, an S-Band
(ST-R), and an S-Band Signal Processor (SSP).
The OMNI antenna will need higher power than the NV OMNIs because of the distance the signal must travel. and demodulates the S-Band signal. decoding, multiplexing, bound signal.
The ST-R modulates
The SSP performs the encoding ,
and demultiplexing
of the in-bound
and out-
The S-Band network receive input and outputs data to
the DMS subsystem.
5.5
Data
Processing
Both the TV and the NV will have isolated Data Processing Subsystems (DPSs).
These subsystems shall support communication,
GNC, tracking, and control and monitoring of the vehicles.
Any
instrumentation
All
formatting
data will
be processed in this subsystem.
and preparing of the data to be transmitted to the Control
Center will be handled by the DPS.
Basically, the DPS constitutes the
"brain" of the two vehicles. The DPS for each vehicle will loaded with
identical
software.
consist of 2 redundant computers
The computers will
art to provide maximum processing capability. commands and data will
All processing of
be conducted by the DPS.
the DPS subsystem are provided in Table 5.12.
55
be state-of-the-
Characteristics of
Table
5.12
Characteristics
O:U
N
Weight
Computers
V
(2)
T
and
and
The
Removal
Debris
track
current than
the
tracking
10 cm
problem, cannot
employed
by
from
Earth.
For
assumption this
as small is
reasonable
requirement
of current
performed
is
additional
tens
The to
are
of
orbital
Established
ground
deployment
of the
debris
to
less
greater
presents
than
assumed
10
that
SPECS,
Inc.
ground
tracking
systems
and
within
change
that
will
believes
that keep
a
cm,
that
system
are
The
in the can
this
will
be
track
of
that the
the
particles[5.23]
will
based DRS
of
them.
This
it was
locate
particles
[5.22].
tracking
biggest
computers
thousands
allows
consideration
The
the
Vehicle
new
under
technology.
Netting track
in LEO.
remove
those
DRS
based
because
upgrading
debris, the
to accurately
it can
in LEO
a ground
as 1 cm
able
NORAD
tracked
be tracked
be
be
before
target
particles
system
must
of our
track
ability
Subsystem
particles
to be
TV
(in^3)
1440
debris
will
Volume
and
100
some
there
NV
35
because
future
the
1440
System
system
for
Power(W)
(kg)
Detection
in diameter
near
meet
orbital
DPS
100
(2)
Tracking
the
35
Computers
v
5.6
of
employ in
a combination
order
radar
will
the
target
56
to estimate
detect
active/passive a rendezvous.
a breakup
trajectory.
Once
to the
guide
the
Transfer
Vehicle semi
is major
Transfer
axis
to
debris,
enter
track
enhanced
active
the
debris.
tracking
For
our
an
active
using
view
the
the
tracking
system,
we
radar
tracking
systems
that
were
considered
infrared,
an
tracking
system.
optical
light
a low
at a distance infrared
shown
of 500
tracking
are
moving
the
system
is able
in
of their
tracking for
Space
Shuttle
is the
LADAR
in
in Space
at
of
[5.24].
system.
This 57
were
an
Ranging)
small
telescope
as
small
as
1 cm
is
the
considered debris
particle that
particles
due
to
are
2 cm
then
determine
how
their
location
considered
and
testing
of
an
Finally, system
scheduled the uses
to third
pulses
in
detects
is
is
by
solar
system
Station
1991
tracking
This
space
field
a
the
to
of
using
system
view
the
Lincoln
particles The
with
MIT's
particles
the
the
narrow
and
system
by
km.
field
their
Detection
and
possibility
Vehicle
tracks
detect
velocity.
debris the
to
size
the
will
approaching
The
Netting
Another
1900
the
of
detect
off
particles
on
can
[5.22].
of
because
sensors,
given
a distance
system
optical
radiation
at
out
performed
sensor
system
direction
our
This
diameter
avoidance
IR
km
is
the
power
Vehicle
possible
ruled
(LAser
detector,
be
(i.e..
from
low
Netting
requirements.
sensor.
This
for
for
that
video
heating.
practical
power
Experiments
the
the
large
or a LADAR
detecting
and
system
using
the
then
torus
detach
vehicle
determine
have
will
the
the
would data.
very
and
After
debris
Then,
debris,
would
the
Vehicle
torus.
of
sizing
their
have
debris
Capture
and
below
Netting
piece
sensors
and
Laboratory
the
trajectory.
the to
kilometers
less),
a rendezvous
distance
of
km
and
sensors
compute
50-100
50
Vehicle
passive
the
established
IR be
to
the
be
collision conducted
tracking of laser
light
to detect the debris and accurately measure its size and distance from the spacecraft.
The LADAR
system is able to resolve the size of
a particle to a microradian at a range accuracy of 0.1 m at 25 km [5.25-5.28]. To
satisfy
Netting
Vehicle,
system.
The
sensor
and
our
SPECS, tracking
an
active
confidence
by
NASA
wide
of
view
field
Netting debris the
vehicle
to
determine
Inc.
in
IR
system the
the size
of
detection to
as
on
our
a combination IR
in Figure
tracking
tracking 5.11.
The
avoidance
system
and
the
select
system
for
our
us
to
will
be
used
Vehicle
the
system
a passive
shown
led
debris,
use
both
collision
Netting
approaching exact
use
system
the
the
decided
will
LADAR
maneuver
the
has
system
This
is
for
it provides
Vehicle. and
requirements
this
initially toward
the
to the
LADAR
debris
and
the
locate
the
debris.
Once
system
will
distance
to
be
used
the
debris.
IR @ 1 900 km v
A
Figure
With
this
debris
and
range
of
data,
5.11
the
Netting
fire
a net
will the
25 km
,.- Ladar@
Tracking
Range
Vehicle and
will
capture
net. 58
Characteristics
continue the
debris
to
close
once
in
on
it is in
the the
The reason that these two systems were chosen was because they each offset the other's weaknesses.
The main disadvantage of
the IR system is that it is unable to determine the size of the debris piece, however it is able to detect the particle far away. disadvantage is that it is difficult
Another
to accurately determine the range
of the particle with the IR system.
On the other hand, the LADAR
system is able to accurately determine the size and the range of the debris piece once it is within
25 kilometers of the Netting Vehicle.
However, it is unable to detect the particles at large distances. Therefore, we have chosen these two systems so that we are able to detect the particles at long ranges and measure the size and distance to great accuracy once the Netting
Vehicle
has
closed
in
on
the
particle. The
tracking
system
will
the
LADAR
system
and
be
mounted
on
propulsion
LADAR about
will
kg.
Most
light
This of
system 100
optical
gives 70
the
collector a total
Watts.
computational
10 Watts
weigh of the used system
These
systems
power.
The
by
the
on-board
maneuvers
can
be
performed
the
location
instead
that
has
to
use
be
of
the
of
ground
transmitted
60
the
IR
module
of
the
about
30
IR
sensor's
to
about
for
focus
weight
interpreted
Finally,
require
of
IR
130
on-board
computers between 59
will the
The
to
the
large
on
the
sensors.
the
with to reduce
will
a power
weight
consumption
amount
will
that
vehicle
will
is due
sensors
computers
systems
system
and
rendezvous
for
IR
a substantial
so
power
Vehicle.
radiation
km
the
computers to
the
of The
Netting
and
require from
system.
weight
the
also data
kg
Watts
have
to
of be
proper the
debris
determine the and
the amount
ground.
particle. particle of
data
5.7
Guidance,
5.7.1
Navigation,
Guidance Guidance
main
tasks.
center
of
mass
determining tasks,
using
and
some The
were
form
of
designed
for using
guideline,
since
this
not
restrictive
consists
of
freedom
gyroscopes.
algorithm, inertial are
three
the
in
information. each
watts
to
and Due
IMU
have
the
IMUs has
to
of power
in
center
satellites
conjunction sufficient
interest
mentioned as using
of
(GPS)
10
drift by
to
an
consume
to
[5.29, to
sources.
a spacecraft,
or
the
Tracking
of
196].
the
Three
IMUs
redundant
for
a typical 47
200].
In either
and
IMU
approximately
inaccuracies,
and
determine
provide
and
of
past
intergrating
requirements
mass
the
two-degree
two
[5.29,
6O
in
and
kilograms
other
vehicle
Each
shuttle,
basic
netting
weight.
above
power
angular
a general
information
the
estimated
updated of
in
the as
with
both
IMU).
reliable
and
In
determined
(or
shuttle
itself
of
velocity,
and
accelerometers
measurement
of the
proven
provide
was
the
two
the
attitude.
generally
vehicle
of
consists or
unit"
aboard
terms
a mass
be
positioning
transfer
vehicle,
periodically
position
the
Furthermore,
rate-gyro,
must
each
are
both
system
second
as position,
measurement
Used
of
such
into
position
orientation,
"inertial
orthogonal
quantities
utilized
inertial
broken
the
the
measurements,
in
IMUs
conveniently
while
interest,
angular
IMUs
overly
spacecraft,
of
be
of determining
spacecraft's
quantities
velocities,
may
consists
of the
Control
Navigation
navigation
first
the
the
and
and The
and
Data
the
IMUs
determining global Relay
Satellite
the
is
System (TDRSS) may be utilized.
In either case, onboard computers
can be used to analyze the time delays and the doppler shifts of radio signals sent to the spacecraft from a ground station through a TDRS.
Given a sufficient number of time delay and doppler shift
measurements (i.e., range and range-rate information),
and given
dynamic models for both the spacecraft and the TDRS, the position and velocity of the spacecraft's center of mass may be calculated. course, it is typically
necessary to provide error modeling,
to dynamic modeling, to filter
out random noise.
Of
in addition
The concept of
using TDRSS for the on-board tracking of near-earth satellites is extensively
discussed by Shank in his article "Automated
Orbit
Determination Using Tracking and Data Relay Satellite (TDRS) Data" [5.30, 1-21]. The decision was made to utilize TDRSS in navigation because TDRSS is also used for the design's communication
purposes.
Further,
onboard computers are anticipated to handle much of the navigation work to minimize ground support. providing
communications
under 5000 km in altitude
Moreover, TDRSS is capable of
and tracking for over 85% of the orbits [5.29, 288].
In addition to the center of mass position information, attitude information updated.
provided
by the IMUs
the
must also be periodically
This updating may be accomplished
by using appropriate
sensors (described below) and an on-board computer.
If the position
vector of the center of mass of a spacecraft is known, it turns out that knowing
the unit vectors to two non-collinear
and Sun at an appropriate
bodies (the Earth
time, for example) uniquely
61
determines
the attitude of the spacecraft [5.31, 140].
These unit vectors may, in
turn, be obtained from Earth, sun, or star sensors. Sun sensors have the advantage that, for near-Earth orbits, inertial
displacement
the
vector from the spacecraft to the sun is virtually
constant over several orbit revolutions,
thereby providing
a direction
that is fixed in inertial space for a time duration of interest (for example, the time it takes to perform [5.29, 155].
an angular momentum change)
A further advantage of sun sensors is that, because of the
sun's brightness,
they tend to be relatively
inexpensive,
reliable,
and
consume small amounts of power [5.29, 155]. Earth sensors generally consist of a scanning mechanism, an optical
system, a radiance detector, and signal processing electronics.
The principal
drawback to Earth sensors is that significant
uncertainties can arise due to the presence of the atmosphere on the horizon [5.29, 167].
However, for near-Earth applications,
they have
the advantage that the Earth is always in view and cannot be confused with
other luminous
sources.
Star sensors are generally the most accurate sensors, but the drawback with these sensors is that they tend to be heavier, more expensive, and consume more power than other sensors. require preprocessed position
They also
data on the star being tracked as well
as extensive star maps and computer software for data reduction [5.29, 186]. Magnetometers are used to detect the direction magnetic field in body-fixed
coordinates.
of the Earth's
Then, knowledge of the
Earth's magnetic field and the position of the center of mass attitude
information.
Magnetometers 62
gives
have the advantage of being
lightweight,
require only a small amount of power, and can operate
through a wide range of temperatures.
However, they often cannot
be used with confidence in determining
the attitude of the spacecraft
because the Earth's magnetic field is poorly [5.29,
known in many regions
180-181]. The criteria
importance,
for choosing the sensors was, in decreasing order of
accuracy, power, weight, and expense.
The importance
placed on the accuracy was due to the extensive docking and debris capture anticipated.
Further, as a result of the accuracy requirement,
magnetometers were not used. all the other three sensors.
Each vehicle, however, makes use of
Even though it requires only two sensors
operating at one time to theoretically orientation, shadow.
three will
determine the spacecraft's
be used for redundancy and for use while in
Also, four star and digital sun sensors will
be aboard so as
to encompass a large field of view, even though only one of each will operate at any given time.
The weight of these sensors and the
power they consume (per vehicle) respectively,
5.7.2
moment
basic gyros
associated
in
CMGs
is
during
the
Control
control
mechanisms
(CMGs)
with
considerable weigh
25 kg and 20W [5.29, 177-190].
Vehicle The
were estimated to be,
CMGs
power. excess that
and
of
is For
600
process
of
RCS
that
they
[5.29,
The to be
some
of
201].
secular 63
vehicles
tend
momentum
cancelling
both
thrusters.
instance, lbs
undesirable
of
primary
large
the
will
and
larger
Another configurations disturbance
be
control
disadvantage consume
CMG
systems
disadvantage
of
invariably torques;
arise as
a
result, CMGs are usually accompanied by an RCS system for periodic momentum dumping [5.29, 200].
However, CMGs offer the capability
for fine tune attitude ajustments, as required in docking retrieval, and they will
and debris
not blow the debris away as an RCS might.
Equally important, if an RCS was used exclusively, the amount of fuel required by the large transfer vehicle possibly years would definitely
limit
shuttle missions, which are relatively
over many months and the mission.
For example,
short, can require over 3600
kg of fuel and oxidizer for its RCS [5.32, 297].
Lastly, based on
representative CMG systems, the CMGs for both the transfer and netting vehicles were estimated to have a mass of 175 kg and to consume 100 W of power [5.29, 200]. A RCS is necessary to supplement the CMGs and provide small adjustments in the position of the center of mass.
The dry weight of
the RCS of the transfer vehicle was roughly estimated using the dry weight of the RCS of the Orbital Maneuvering Vehicle (OMV) as a guide, because both vehicles perform similar roughly the same mass.
tasks and are of
The dry weight RCS estimates for the netting
vehicle were obtained by scaling the dry weight RCS estimates of the transfer vehicle down to 25%.
The OMV RCS consists of 28 hydrazine
thrusters weighing 5.45 kg apeice and with a thrust of 15 lbs [5.33, 30, Appendix
1].
The RCS fuel requirements were difficult
to estimate because,
as of now, it is not known exactly how large a role the RCS will play in relation to the CMGs. providing
virtually
It is anticipated that with the CMGs
all the attitude control and with the possible aid
of the ion engines for fine-tuning
the position of the center of mass, 64
the role of the RCS will be minimized.
For calculation purposes,
upper limits for the combined fuel and oxidizer masses for the Transfer and Netting Vehicles were speculated to be 1500 kg and 400 kg, respectively. subsystem with
A summary of each component of the GNC
its corresponding
weight, power, and volume
estimates is given in the Table 5.13.
(The volume of the RCS systems
include fuel volume estimates based on the bulk density of hydrazine
and nitrous oxide being 1200 kg/m3.)
Table
5.13
NV
Mass
and
TV
Weight
(kg)
Power
and
Power
(W)
for
Volume
Sensors
25
20
.5
IMUs
3 0
140
1.0
CMGs
17 5
10 0
1.0
*****
TV-4.6,
RCS
(dry)
Total
TV
(dry)
3 95
26 0
7.1
Total
NV
(dry)
27 1
260
4.0
*
TV-165,
Includes
In
fuel
addition
moments
of
Vehicle done
will
be
because,
pointed
to
inertia
one in
active the
designed
the
on rev
(m3)
NV-1.5
*
estimates.
the and
while
approximately be
volume
NV-41
GNC
nominal
for the per
appropriate
control
of
gradient
transfer
orbit, period
direction 65
mentioned
orientation
gravity
orbital
systems
the
above, Transfer
stabilization. the
so at all
vehicle
that times.
the
the
This must
is
spin
ion
engines
(The
ion
at can
engines
do not provide
ideal delta v's; but rather, operate continuously
throughout the transfer.)
A spin rate of one revolution
period is ideal for gravity gradient stabilization. stabilization than the
the yaw
moment
moment
principle
Finally,
of
was
[5.31,
inertia
idea
which
are
to
the
Vehicle.
The
requirements
for
spinners
deals
bodies
[5.31,
175-188].
would
correspondingly
would
restrict in
terms
the
structure
idea
that
four
Newtons
light
and
stresses the
the
dual
Further,
the
require design
of
the
extra
of
the
Transfer
ion of
thrust.
resulting
by
vibrations
Vehicle
very
flexible.
fast-spinning
would
be
66
stability
extra
mass
vehicles
into
space
transfer
structure
Vehicle
this
orbit.
mainly
would
unacceptable.
using
axisymmetric
designed
only
Finally, with
provide
as a whole
would
mass,
on
and the
It
axis.)
Vehicle
Transfer
spinner,
was
the
the
the
the
non-axisymmetric
with
getting
than
that
Transfer
and
of
for
greater
three
available
together
Therefore,
a huge,
of
required
working
correspondingly induced
terms
fuel
engines
size
a large
in
size
largely
the
these
the
theory
be
for this
greater
is assuming
about of
be
should
(This
control
due
should
turn
alligned
of passive
rejected
in
203].
Transfer
and
the
inertia
inertia of
the
spinner
nature
of
of
moments
a dual
The criteria
is that the pitch moment of inertia
roll
per orbital
and
not even
about
is quite
withstand if it
the
the could,
SUN EARTH
DOCKING
& STAR TDRSS
INFORMATION DATA
SENSOR
REMOTE CONTROLS J COMPUTER I
CONTROL CALIBRATION
v_
'_
J
y"-
SYSTEM CONTAINING SPACECRAFT DYNAMICS
COMPUTER
RCS/CMG
IMU
Figure
5.8
Netting The
launching
material. simple the
netting
subsystem the
is
Integration
composed
retrieval
will
be
made
from
The
nets
will
be
spinning
to
spring open
the
system, net
of
system,
nets
compressed
perimater
GNC
System
Subsystem
system,
The
5.12
Kevlar, when and
with
67
and
there
centrifugal
four
parts:
the
storage
a high they
be
forces.
nets,
the
volume.
strength are
will
the
composite
launched four A
by
masses Kevlar
a on net
1
meter (m) in diameter, 1 millimeter
(mm) thick and with four 0.23
kg masses on the perimeter will have a mass of 1.92 kg.
2.5 kg was
used to include an extra amount of mass for the launching system. 2 meter diameter net and launching system will kg.
A
have a mass of 6.0
A 3 meter diameter net and launching system will
have a mass
of 11.5 kg. After the net has captured the debris (see Section 6.2 for more details on launching the net and capturing the debris), the net and debris will
be retrieved
the Netting Module. Netting Module. approximately
by a tether connected between the net and
The tether will be wound up by a winch in the
The netting winch should have a mass of
50 kg, a volume of 0.0063 m3, and a power
requirement of 78 W (based on small automobile winch as a model). There will
be only one winch per Netting Module, with a separate
cable for each net.
These cables will be able to be deployed, braked,
and retrieved independently.
The mass of the cables is expected to
be no more than 16 kg (calculations based on 20 steel cables 2 mm in diameter and 100 m long). The sum of the cross-sectional areas of the storage volumes will not exceed 75% of the area on the front face of the Netting Module in order to ensure structural rigidity. volumes
Three sizes of storage
were considered: • A 20 cm diameter, 50 cm long cylinder Could safely hold a plate 14cm x 14cm or smaller Would use a 1 m diameter net A 40 cm diameter, 60 cm long cylinder Could safely hold a plate 28cm x 28cm or smaller Would use a 2 m diameter net
68
A 90 cm diameter, 110 cm long cylinder Could safely hold a plate 63cm x 63cm or smaller Would use a 3 m diameter net
The dimensions of a plate that could safely fit in each cylinder was taken by assuming that the greatest possible length that could fit across the cylinder would be a plate with a length the size of the diameter.
The length of the sides were chosen by considering the
worst case: the plate could be turned so that its diagonal is being pulled across the cylinder.
The sizes for safety are therefore the
diameter of the cylinder divided by the square root of 2. Furthermore,
three different
Netting
Module
configurations
were examined: •
NM20 - has 75 20cm holes Total Storage Volume - 0.94 m3 Mass of nets and launching systems - 187.5 kg NM20/40 - has 18 20cm holes, 9 40cm holes Total Storage Volume - 0.9 m3 Mass of nets and launching systems - 99 kg NM20/40/90
has 12 20cm holes, 6 40cm 1
90cm
holes,
holes
Total Storage Volume - 1.3 m 3 Mass of nets and launching systems
5.9
will
Structural
Materials
The
Module,
Propulsion
aluminum,
a proven
be
Composites expensive
Netting
made
of were
for
our
considered, system.
but An
they estimate
69
Module,
and
material were of
in
space
judged the
Transfer
to
structural
Vehicle
flights. be
too mass
was
- 77.5
kg
made by assuming that each of the vehicles was a cylinder both ends with a skin thickness of 2 centimeters.
closed at
A 10% factor was
added to this figure to take into account the internal support structure. The
subsystems
allocation of
volume)
volume
for
diameter
located
at the
hollow
cylinder
thickness Netting
of
meter
mass
The
will
engines
a structural Since
navigation
on
be
mass this board,
0.042
of will
shielding
need
so
a 3
meter
subsystems
2 meter
9.163
diameter
to
a
and
a skin
unloaded kg.
m 3 of space.
will
space
diameter
It will
provide
1933.9
kg.
not
a manned
be will will
from
be
needs 0.2
no
and
and
70
21.021
cylinder
satisfy
A
this
will
have
a
to
be
mission
to
for
this
heavy
cosmic
the
a 3 meter
large
space
and
protect
m 3 of
with
m 3 extra
need
suffice solar
contain
long
(the
paint
be
kg. will
there
to
the
a 2
collar,
1653
to have
m 3 extra
Vehicle
used
assumed
approximately
meter
with
docking
of
(m 3)
volume
cylinder
dimensions,
need
a 2
provide
1466.8
inside).
will
with
components,
diameter
Radiation
of
Transfer
subsystem
be
will
of
meters
subsystem
The
a diameter
these
a mass
cylinder
It
with
Module
long
was
a fuel
cubic
of
meter
Module,
With
have
7.91
requirement.
long
centimeters.
(including
a summary
A 2.52
Netting
meter
will
for
this
the
Module
approximately
5.14
satisfy
0.5
long
structural
will
vehicles).
Propulsion
requirement.
ten
all
of
Netting
require
Table
back
4
the
(see
will
Module The
2.93
will
each
requirements meter
for
no
and
nuclear
radiation computer
radiation.
to
fit
will
the have
reactor shielding. and
Table
5.14
Summary
of
NM
Requirements Vehicle
(m 3)
Configuration
20
Subsystem
Volume
20140
TV
PM
20/40/90
$$)I($$$
Structure Netting
0.91
0.95
1.31
Propulsion
******
3.450
8.200
Power
******
0.0121
0.72
0.850
2.750 *******
Thermal
3.95
3.95
3.95
Tracking
******
******
******
0.7802
Comm.
******
******
******
0.047
0.086
GNC
******
4.000
7.100
DPS
******
0.024
0.024
Fuel Total
1 space
for batteries
3.05
2.65
7.91
7.91
7.91
2.140
3.01
2 1.02 1
9.163
only
2 space for sun. star.and earth sensors only; LADAR and IR sensors are mounted on body
However, will
be
impact
in
3M
near
shielding
will
[5.34].
than
be
see 3
has
if it
km/s
Netting
been
as
have
Nextel tested
stop
and
the
concentrations We
called
would
known
Vehicle
dense
needed.
fabric
Nextel
to
the
relatively
ceramic
engineers higher
because
or
lightweight by
1l
by
particles
hypervelocities.
71
of
decided that
Transfer
is
debris,
to
use
being
Johnson
A
debris a new,
manufactured
Space
travelling
Vehicle
at shield
Center velocities composed
of
4 layers of Nextel and a thin aluminum plate has successfully stopped a 1 cm sphere of aluminum travelling
at hypervelocities
[structures. 1]. The debris shield will
be composed of 4 sheets of Nextel, each
with a surface density of 0.123 g/cm 2 (4.92 kg/m 2) The sheets will
[structures.l].
have to be spaced three inches apart and the skin of
the spacecraft will
take the place of the aluminum plate (the plate in
the NASA test was 80 mil, or 0.203 cm thick).
This shield should
stop particles with a diameter less than 1 cm, the small debris our system is not targeting. The mass of the shielding required to cover the front of the Netting Module and the perimeters of the Propulsion Module,
and Transfer Vehicle
is approximately
Module,
359.6 kg.
Netting
This
includes a 10% overestimate to take into account the structure that will
be needed to support the sheets of Nextel.
5.10
Fuel
Requirements
The masses of the other subsystems, as well as their volumes, played an important role in the calculation of the mass of the fuel needed.
The calculations
used the ideal rocket sizing equation
mass of fuel = (mass of spacecraft)x(1
e-dv/g*Isp)
where dv = velocity
change required to change spacecraft's orbit
g = the acceleration due to gravity Isp = the specific impulse of the fuel and the following
assumptions
72
Mass of Netting Module is NM20 2,432.5kg NM20/40 - 2,350.4 kg NM20/40/90 2,332.6 Mass of Propulsion Module is 2672.0 kg The Netting Module completely fills its nets with maximum size debris for each hole (masses for 2 cm thick aluminum plates) 14cm x 14cm plate - mass of 1.06 kg 28cm x 28cm plate - mass of 4.23 kg 63cm x 63cm plate - mass of 21.43 kg Fuel is Hydrazine-Nitrous Oxide mixture Isp = 318 seconds density = 1200 kg/m3 [5.35] The delta v needed to capture each piece of debris (Data obtained from Himawari 1 rocket booster breakup in July 1977. See Appendix C) delta v = 15 m/s The Netting Vehicle collects the smallest pieces of debris first, then moves to larger pieces
A program (a listing is included as Appendix D) was written to iterate the amount of fuel needed for each of the Netting Module configurations
to collect all the debris they can hold.
The program
added an extra 10% at the end to take into account proximity operations when capturing the debris.
The NM20 configuration
would require 2,972 kg of fuel and 2.48m3 of storage space. The NM20/40
configuration
requires 796.1 kg to perform
its mission, and
the fuel will
take up a volume of 0.66 m3.
configuration
required 552.2 kg of fuel and 0.46 m3 of volume.
73
The NM20/40/90
All structure reduced.
three configurations length
can therefore be used, although the
of the 20/40 and 20/40/90
configurations
To maintain an extra volume of approximately
lengths of the NM20/40
and NM20/40/90
can be
0.5 m3, the
structure can be reduced
to 2.22 m, reducing the structural mass by 102 kg. Similar
calculations
were performed
to calculate the fuel
needed for the fully loaded Debris Removal System to go from the Space Station to the parking orbit. We included a 30 degrees wedge angle or 30 degrees inclination
change. The total fuel mass needed
for the Transfer Vehicle was 3400 kg. A complete summary of the vehicle masses, using this new data, is included as Table 5.15. Table
5.15
Summary (All
74
values
of in
Vehicle kg)
Masses
Vehicle NM 20
20140
1653.o
1551.o
Subsystem Structure Netting
Configuration
165.o
253.5
20140190
1551.o
PM 1466.8
TV 1933.9
143.5
******
******
Power
******
******
436.0
436.0
235.0
303.0
Tracking
******
******
130.0
*******
Comm.
******
******
80.0
140.0
GNC
******
******
27 1.0
1076.03
DPS
******
******
9O.O
79.0
79.0
I 16.6
2972.0
796.1
552.2
170.0
Thermal
436.o
Shielding Fuel
180.0
255o.o
Propulsion
23.4 .i
1453.0.2
mass inclu(ledin GNC, tracking,& comm. 153.0 3400.0
Total
- Dry
2432.5
2231.0
2209.5
2502.0
7608.9
Total
- Fueled
5404.5
3027.1
2761.7
2672.0
I 1008.9
assumes
1995 technology
i Battery13 kg Solar Array - 10.4 kg (assumes
35Vo etTeiciency)
2 Battery741 kg Solar Array712 kg (assumes
55% efficiency)
3 includes
mass of" RCS fuel
75
6.0
System
6.1
Debris
Integration Removal
The
final
Module,
Netting
6.1,
and
6.2,
transported
later
into
Module,
and
pre-launch
After
all
be
to
consist
of three
a system to
collect
of debris
Earth
in be
flight,
shown two
65 have
been
space
NM20 it
may of
be
in Figure
6.4.
is
shuttle
(the
6,750
The
total and
22,231.4
the
The
for
Propulsion
modules
debris
kg.
used
debris. one
collected,
approximately
be
Modules,
of
18,637.1 the
the
NM20/40/90
pieces
in Figures
can
amount
Netting
with
shown
However,
as
Propulsion
system
a greater
Vehicle,
will
mass
mass
that
three
kg.
of
the
will
need
unfueled
kg.
Retrieval important
where to
it is. estimate
than
2000
is
within
when
it
these
sensors, as
collect
the
are
complete
initially.
approximately
most
sensor
used
it can
of
Modules)
knowing
be
Transfer
be
Debris
debris
shuttle
pieces
returned
is less
one
module
the
The
IR
with
the
will
Netting
6.2
space
mass
system to
the
not
of
Vehicle
that
will
NM20/40
Transfer
order
since
DRS
and
configurations
In
will
The
and
Module,
missions
one
dimensions
6.3.
configuration
System
the
possible
part The
in
capturing
Netting
Vehicle
will
for
particle
a trajectory
kilometers, 25
later
kilometers.
Netting to
and
Using
Vehicle
facillitate
will capture.
76
the it
a piece
of
first
the
onboard
when
the
distance
LADAR
sensor
use
the
information to
is
use
will
attempt
debris
the
get
derived as
close
from to
the
Top
View
Side
View
2
Body
Mounted
Solar
Aray
2.93
IR LADAR
Sensor
Sensor 1.92 Fuel/Electrical Connection
OMNI ttenna
(I of 3)
Docking Trunion
All dimensions
Front
View Figure
6.1
77
Propulsion
Module
in meters
Top ,A
View k,
2
_1.92
i,
!
!
i
I
i
o .L___
I L
0-5 ,qP
All dimensions
in meters
Fuel/Electrica] ' Connection
Docking Trunion
Netting Holes
Front
Figure
Trunion Lock
Rear
View
6.2
Netting
Module
78
(NM20/40/90
View
Configuration)
All dimensions
in meters
m
Docking Bracket
0-5
:
]
____U
Communications Antenna
1_
illllllUlilili iiliilliliilll fliUillllUilli| IllliilllOlli|
'l
tmtit2 Front
View
Rear
View Top
View
6O
Engine
Figure
6.3
Transfer 79
Vehicle
Nozzles
iiiiiiniiiiiiiu iillilUiliiliil llniiliililinil lillilliUilliil Illllllllllllll Illllllllllllll IIIIIIIIlllllll Illllllllllllll Illllllllllllll Illllllllllllll Illllllllllllll lllllllllllllll IIIIIIIllllllll IIIIIIIIIIlllll IIIIIIIIIIIIIII IIIIIIIIIIIIIII Illllllllllllll
Ullllllllllllll IIIIIIIIIIIIIII IIIIIIIIIIIIIII Illllllllllllll Illllllllllllll
NM
NM
TV
IIIllll
NM liiliil
ilnllllllllllll iiiiiiiiiiiiiii
mmmmmmmmmmmmmmm mmmmmmmmmmmmmmm IIIIIIIIIIIIIII IIIIIIIIIIIIIII Inillllllllllll Illllllllllllll Illllllllllllll Illllllllllllll
lllllllll llllllllllllll llllllllllllll
mmmmmmmmmmmmmmm mmmmmmmmmmmmmmm Illllllllllllll Illllllllllllll lUlllllllllllll Illllllllllllll Illllllllllllll lUUllllllllllll Inlllllllllllll imlllllllllllll Illllllllllllll IIIIIllllllllll
Figure
6.4
Debris 80
Removal
System
Figure
While the
the
sensor
This
is
reeled
This
into
the
size
because
diagonal
that still
safer one
to of
store the
cm
in
by
a tether.
on
the
on
hole
the
50
cm
would
of this
launch and
The
is spun
when
perimeter (assumed
will to
open be
10
it. cm
is
the
size
of
netted,
debris
when cm
to in
be
14.5
crn.
cm
a longer in.
The
it would
Therefore,
a net
at the
debris.
to
the 6.5.
the so
20
The
Netting that
is generated and
debris.
will
20
but
it is launched
diameter
the
in Figure
connected
in
be
it is pulled
cylinders,
the
example,
with
launched
spin
it
For
rectangular
This
81
interpreting
storage
is shown
system
also
cylinders.
be
a spring net
20
it is
storage.
for
be
Net
is
of the
limit
of the
cylinders
debris
could
of the
this,
for
size
upper
one
one
dynamics
by
cylinder
50
the
the
impinge
it in
approximate
after
debris
fit
to
estimate
the
could
doing
Module
the
the
is
Netting
above
might
The launched
the
is just
order
since
sensors
cylinder
debris
in
important,
back
Deploying
Vehicle
information
very
suppose
Netting
6.5
in cm
the the long)
be from
net Module masses
launch because
is
the
end
like
masses
rifling
are
in
a
in
gun
slots
that
spiral
along
the
length
of
the
tube,
barrel.
Perimeter
Mass
Compressed Spinning
Grooves
Figure
This cylinder
1.6
launch and
grooves
in
hitting open which
the
at
inside
some
1 to
cylinder
is the
debris
is The
activated
by
a
small
5
shown
of
1.1
walls
of
net
hit
contained net
will
storage
from
the
braking accelerometer
the
back
of
perimeter 6.6.
If
leaves
there
the
no
will
net
tube
is
and
problem
cylinder. Netting
storage
mases the
the is
each
is
net
Module,
in
spinning
at
travelling
with
The
fit
the
will
masses fully
depending
on
from. the
inside be
it
Tube
the
the
Figure
m/sec,
launched has
at
and
when
meters is
located
in
second
Launching
6.6
midline,
as
speed
the
Module.
(by
wall
a
is
the
per
Once the
cylinder
along
revolutions
forward
Net
debris, when
closed
the
with
tether. on
the
net
will is
be
closed
reeled
pulley
When
a
has
net),
or
82
collision when
the
so
into
mechanical
the
a
net
net
the that
been has
that Netting is
detected reached
the
end
of its
perimeter
will
tether like
is
tether,
the
continue
to
redistributed
will
be
braked.
move
forward
via
pulleys
and
has
been
captured
The
until
masses
the
cables
to
on
tension
pull
in
them
the
the
together
a cinch. After
net
will
the
debris
be
reeled
Vehicle
will
not
relative
velocities,
used
to
does
not
rotate
back be
the
wrap
6.3
into
able
the
to
and
storage
approach
the
moment
gyros
Netting
Vehicle
during
the
contained
cylinder.
control
around
without
RCS
the
a net,
Because
debris and
in
the some
thrusters
retrieval
the
Netting small
will
so
that
be the
net
vehicle.
Docking In order
Propulsion Netting
for
this
Module Modules
will
will
have
between
the
Propulsion
establish
trunions
on
the
flight
vehicle
be
similar,
but
fluid
Propulsion
will to
to
Module
fuel
of (see
have
6.7).
since
connections
couplings
not
or Netting
couplings. 83
docking
the
be
and
all
docking will
uses to
electrical
the
will
Module
Vehicle
Our
between
storage, The
module
fuel
they
for
the
have
to
interfaces.
Maneuvering
Figure
the Modules,
Vehicle.
power
propulsion
All
other
Netting
its
work,
Netting
each
the
for
to
the
Transfer
and
umbilicals
transfer
Module.
Module,
and
Orbital
[6.1]
Propulsion Netting
the
with
with
the
Module for
System
dock
with
perimeter
connection
to
to dock
to dock
proposed the
Removal
have
have
connections NASA's
Debris
will
of them
the
tether
connections Modules
be
needed
Module
connect
controlled
for to
with
mechanism
Netting will
four
Netting
Transfer
will and
a
and by
the
Module Vehicle
Lock
Docking
Figure
7.0
Debris In
this
debris.
in
spring
[7.1,60-72]
prevention that prevent costly
section
design refer
section techniques
modification orbital mission
debris
discuss
of
to
part
this
the
therefore design
mission is
active
far
concepts
discussion
report
and
of
for
will
a detailed
will
Mechanism
Concepts
we
Since
we
This
Docking
Prevention
orbital the
6.7
Trunion
orbital
of
the
contain
more
84
of
this
topic
included
was
working
group
overview
of
debris
to
the
report. a short
and
economical removal.
prevention
debris
alterations.
hardeware
debris
on
for
This space than
relates practices a complex
to and
fact
7.1
Self
Disposal
By Earth
deorbiting
escape
or
be
active
7.1.1
by
payloads
can
without
orbiting
below
meters
can
years
to
of
ballon
800
kilometers, the
of
reamain
inert
for
inflated
after
a rocket
or
the
drag
device
does
not
attitude
7.1.2 Solar high
part
orbits.
in
to
by
a series
of
to is
an
are
maintain needed
option
or
years.
for
engines. orbit
escape
85
of
its
it is
any
specific
be
to
safely
ballon
mission.
could
The
simple,
15
several
would
have
that
about
from
The
orientation
of objects
passive
system
and and
into
trajectories.
might higher
in
and be orbits
However,
be
main
passive,
disposal
They
the
objects
device would
of
[7.2,4-5].
a relatively
geosynchronous Earth
and
increased area
For
satellite
deorbit
is
be
a diameter
the
completes
concept
storage
onto
of
many
satellite
be
sails
propellant
satellites
or
effective
mass.
with
payload
of up
need
might
Solar
satellites the
a period
Sails
moving
orbits
space
can
the
its
proposed
mission
Solar
no
send
the
system
require
achieved
increases
lifetime
This
control
sails
the
a satellite
a ballon
orbital
as
no
higher
of
be
on
increasing
included
satellite
can
drag which
weeks.
the
into
methods.
atmospheric
significantly
reduce
of
This
and
several
advantage
them
contamination
prevented.
a large
satellite
inserting
Devices
effect
deploying
or
further
devices
Drag The
Spacecraft
trajectories,
environment passive
of
very
they
used or
for to
deployment and control of the solar sail might present significant technical
challenges.
7.1.3
Deorbit
Engine
Another method for self-disposal is the addition of a seperate system for deorbit at the end of the operational lifetime. with a conventional be effective
propulsion
Deorbit
system is an approach which
for all orbital altitudes (for circular
would
orbits above 25,000
kilometers, an escape from Earth orbit is less costly than a deorbit maneuver). wight,
Such
but
is
altitudes mass
is still
below
using
its
to
propulsion
satellites
propulsion
system
drag
raising.
Adding
be
for
the
station
keeping
motors
to
useful
life
of
spacecraft
has
ended.
efficient.
This
agencies of
a satellite
satellite into
for
or
other
policy
their the
several
a "Graveyard
active
payload
retrieval.
appear
has
remaining hundred Orbit"
a small
devices already
geostationary
designed
a controlled
enable
engines
the
to
be
For
a lower-
[7.3,5].
can
or
additional
than
devices
impact
the
increase
expensive
packages
and
orbit
naturally
Fuel
stages
own
less
kilometers
Additional Upper
would
much
700
alternative
7.1.4
a system
act
been
which
does
86
ocean
as
engines
once
deorbit
is
method
above
requires
therefore
adopted
not
and
would
keeping
kilometers
deorbit of fuel
satellites. station
self-disposal
percentage
This
and
for
At
the
fuel
a number
end is
of
used
to
with
cost of
the
geostationary interferre
no
relatively
by
space
lifetime
boost altitude
the
the
the
geostationary
ring,
collision
thereby reducing the
probability
significantly.
7.2
Subsystem By
Redesign
modifying
current
the
production
of
By
minimizing
the
redesign
and
removal
procedures
7.2.1
additional risk
mission
upper depletion
of
can
pressures.
all
i.e.,
fuel
and
oxidizer
the
tanks
due
contractions
main
be
components
widely
prevented.
breakups
by
extremely
hardware
costly
active
reduced.
orbital
debris
design
pressurized
hold
as
orbital
alterations,
to
Therefore,
standard,
can
and
Redesign
One
of
subsystems
debris
future
be
contributors
stages.
space
design
Rocket Main
spacecraft
change
as
engine
possible
to
structural
the
vehicle
is
propellants
is
long
fatigue goes
arrangement
should
enough
vented
from
to the
(repeated
in and
out
breakups
reduction
restarts
on
been
the
and
experimental the
have
for
of be
assure
the
gas made that
tanks.
as
much
Leakage
expansions of the
of
of
and
eclipse)
has
to
be
considered[7.3]. 7.2.2
Seperation Currently
rockets
stages
separation
launch
they
use
and
the
mechanical Space
most
because
rocket
Center,
Mechanism vehicles
release Houston
are
explosive
payloads.
related
Redesign
In
stage
order
mechanisms system
is
referred
need
currently
[7.4]. 87
to
to
connecting provide to
be
being
as
"dirty"
bolts
a clean
to stage
redesigned. developed
separate
Such at
Johnson
a
7.2.3
Increased The
systems with
design
needs
are
space
when to
critical
change:
The
fail.
Reusable
soon
be
expendable
risks
as
are
be
and
stages
high
replaced
which
by
needed is
litter
the
or
beginning
multi-purpose
periodically
transfer
of
elements
longer
philosophy
and
cost
spacecraft
no
upgraded
space
associated
of the
and
they
future
costs
because
maneuvering
upper
of
and
vehicle
could
repaired
orbital
design
"expendable"
satellites
can
the
the
launch
systems
design!). the
all as
which
Hardeware
Generally,
or abandoned
single-use
in
account
hazard.
hardware,
platforms
replace
Reusable
applied
into
debris
jettisoned
of
philosophy
to take
a growing
launching
Use
(modular
vehicles
could
orbital
environment.
7.2.4
Improved Advanced
design
shielding
can
greatly
by
meteorite
and
as
a
of
secondary
would
7.2.5 Another paints
and
could
reduce
concepts
minimize space
multi-layer
then
Shielding
debris act
as
the
debris
bumper
debris
Redesign
of
main
source
degradation
creation
by
significantly
the
of
than
debris
Alternative elements
88
structure
reduce
All
the
shielded
debris
caused such amount
surfaces
"sources".
Coating
orbital
those
debris
multi-wall
impact. rather
spacecraft
secondary
can
Protective
coatings.
of
future
A
"sinks",
of
to
impacts.
system
created
protective
applied
is
microparticles
durable by
atomic
bonding oxygen
from agents and
the
harsh
coating
thermal to
effects
Management
8.1
Management The
Program
leaders
into
of
Figure
cause
8.1
SPECS,
paint
shows
Inc. and
structure
effort
fall
realistic
The
Technical a common
and
oversees
involvement.
developing
the
and
combines
three
designed
a diagram
also
the
Program
works
long
term
contact
and
distributed
with design
and
point
of
the
the
of
a
subgroup to
facilitate
the
complete
goals
that
the
the The
subgroup
project
decisions
and
design design the
Weekly by
the
the
between
group.
three
with
goals
coordinates
of
Manager's
closely
Manager
within
aspects
administrative
Manager.
combined
directly
into
all
The
Program
in communication
intermediate
to
Manager,
support
Manager
Manager
collected,
by
a Technical
process.
Program
leaders
adopted
organizational
Program
level
provides
not
structure.
coordination
work
structure
Manager,
engineering
The
on
order
Structure
an
management
The
in
Proposal
management
general
high
space,
fleck.
8.0
the
in
status
leaders
progress
responsibility. Technical
to
toward
the
Manager
milestones. effort three
and
subgroup
reports
Technical Technical
at a
are
Manager Manager
to
aid
must
develop long
term
milestones. The engineer's into
subgroup design
a workable
leaders
are
philosophy product.
responsible
and The
integrating
subgroup 89
for
leaders
directing
each
the
individual's
provide
a means
effort of
communication
between
information
is
the
separate
subgroups
when
cross-
required. Group Leader Erika Carlson
I Technical
Lead
Foley Weems
r Mission Design Don Chambers
Mission Support Steve Casali
• Structures • Propulsion • Environment Andrew
Lalich
Garner
Geisler
Manfred
Mission
Operations
Manfred
Leipold
• Trajectory • Control
• Ground Support • Communication
• Monitoring
• Maintainability
• Data Processing System
• Budgeting • Mission Scenarios
Leipold
John Parry
Don Chambers
Foley Weems
Richard
John Parry Garner Geisler
Mach
Erika Carlson
Figure SPECS, responsible
8.1 Inc.
for
responsibilities. members
is
SPECS,
consists
the
of nine
engineering
As
a result,
facilitated.
Most
more
subgroups.
Any
arise
are
transmitted
quickly
Inc.
Organization members
tasks
and
communication of
problems to
that the
Structure are
management between
the
group
members
or
requests
for
the
management
subgroups.
9O
dually
the belong
information and
the
group to two that other
or
8.2
Subgroup The
effort
organizational
into
Mission
Responsibilities
three
structure
subgroups:
Operations.
aspects
of
the
The
Mission
mechanical and
systems
and
Inc.
Design,
Mission
subgroup
Design
design any
SPECS,
divides
concentrates
the
design
Support
on
and
particular
project. subgroup
development
research
Mission
Each
overall
of
of
focuses
of
the
the
propulsion,
robotic
primary
on
and
the
structural
secondary
designs.
environmental,
development
is
the
and
and
All
electrical
responsibility
of
this
subgroup. The
Mission
affecting
the
dynamics
and
design
any
must
mission
this
planning
Operations
its
operation.
the
vehicles
area.
the
support
in
this
mission
scenarios for
the
this
handles
by
subgroup. the
design
mission
maintainability,
also
the
monitoring
by
required
and
area.
commanding,
identified
Communication, are
aspects
analysis
developed
are
considerations
critical
Trajectory
systems,
develops
ground
these
is
safety, the
and
Mission
team.
Task A
handles
requirements
Any
in
team
processing
Operations
perform.
developed
of
data
instrumentation Mission
8.3
and
control
Additionally, and
Support
Development
project
design
effort
Figure
8.2
design
process
timeline was
that
developed
illustrates were
the
displays to
aid
project
identified
the in
schedule. to
help 91
major
meeting The control
milestones the
project
critical the
of
the
deadline.
paths
development
of
the of
the
project.
Figure 8.3 depicts the PERT/CPM critical
path chart.lank
page for timeline Figure 8.4 describes the problem solving method SPECs, Inc. employs.
Problems are detected by an individual
evaluated according to criticality. internal to the subgroup. subgroup level.
Minor
or a subgroup and
problems will
Research on the item will
Again, the item will
entire group must become involved.
be solved
proceed at the
evaluated to determine if the The item can either be
discussed and solved at the subgroup level, with a presentation of the solution to the full working
group for education, or the item can
be referred to the full group for a discussion and solution.
8.4
Workload
Considerations
Because of the size of SPECS, Inc., each engineer is involved in several tasks.
To keep track of individual
workloads,
manpower
utilization
charts are collected and updated weekly by the Project
Manager.
As an estimate of the total man-hours required for the
project, it is assumed each engineer devote 12 hours a week toward the project,
and each manager contributes
92
15 hours weekly.
Blank page for timeline
93
Blank page for Pert/CPM
94
Figure project.
8.5
displays
The
1722
man-hours.
hours
to guard
total
the
resulting
effort
required
This against
estimate over
and
I Detect
manpower for will
estimate
for
the completion be compared
under
working
the
total
of the project to the
the
actual
is
man-
engineers.
Problem
I Alert Subgroup
I AlertGrOup
I Rese+a IProblemc
_
Discuss and Solve
Discuss and Solve
_I Figure
8.4
Solution Present
Problem
Solving
95
14 with
SPECS,
Inc.
200O
m L= :3 o
"|
1000
tm
0
5
10 Week
Figure
9.0
Cost
9.1
Personnel
Pay follows" Lead, consultants,
8.5
Manpower
15
Number
Estimates
for
SPECS,
Inc.
Proposal Cost
scales
were
Engineers, $22.00/hr;
Estimate
derived
from
$17.00/hr; project
the
Request
Sub-Leaders, manager,
$75.00/hr.
96
$25.00/hr;
for
$20.00/hr; and
Proposal Technical technical
as
Table Weekly
9.1
manager
1 technical sub
9
engineers
5
hours
Projected
leaders
TOTAL
330.00
$22/hr:
720.00
$20/hr:
1530.00
$17/hr
375,00
14
cost
$ 3330.00
estimate:
$
weeks:
$
and
material to
furnished
equipment
mainframe table
46620.00
4662,00
ESTIMATE
Material
the
375.00
estimate
expenses
in
Costs
$25/hr:
personnel
for
error
The
and
Projected
consulting
weekly
10%
@
@ @
of
cost
@
lead
3
total
9.2
of
breakdown 1 project
plus
Formulation
date
Hardware and
and
(GFE)
computer
Costs
hardware
those
51282.00
of
cost
previous
consists time.
A
below
97
of table
estimates design computer of
are
groups.
based Government
hardware,
anticipated
on
costs
software, follows
Table
9.2
Anticipated
Hardware
Costs PROPOSED
Macintosh
software
IBM
PC-AT
CDC
computer
modeling
software
$
peripherals: and
mainframe
of
photocopies
and
500.00
peripherals:
50.00
time:
200.00
design: @
35.00
$.05/each"
transparencies
@
miscellaneous
supplies:
70.00
$.70/each:
80.00 $
SUBTOTAL plus
10%
Total
error
$
TOTAL
3235.00 323.50
estimate
Estimate
ESTIMATED
23O0.0O
3558.50
COST PROPOSED
personnel material
and
GRAND
TOTAL
COST
TO
1.1
1.3
51282.00 3558.50
cost $
$
(12/3/90)
54,840.50
35,756.13
References 1. Baker, Nijhoff
1.2
hardware
DATE
10.0 Section
$
cost:
Report National
Howard
A.,
Publishers"
Texas,
Boston,
on
Orbital Security
"A Short Course Orbital Debris", Mar
Space
Debris, Council, on Dealing Southwest
19-22,
Debris:
Policy
and
Massachusetts, Group D.C.,
with the Research
Growing Institute,
98
Martinus
1989.
Interagency Washington,
1990.
Law.
(Space) February
for 1989.
Challenge of San Antonio,
1.4
McKnight, D. S., Chobotov, V. A., "Artificial Updates Portland,
1.5
Chobotov, Resulting Rescue 223-241
Section
2.1
2.2
Space
and Insights", AIAA Astrodynamics Oregon, 18-19 August, 1990. V. A.,
"Dynamics
Collision 1986-1987,
of Orbiting
Hazard Science
Debris:
Conference,
Debris
to Spacecraft", and technology
Clouds
and
Space safety Series, Vol.
and 70,
pp.
of Orbiting Debris Clouds and to Spacecraft", Space safety and and technology Series, Vol. 70,
pp.
2.0
Chobotov, Resulting Rescue 223-241
V. A., "Dynamics Collision Hazard 1986-1987, Science
McKnight, Updates Portland,
D. S., Chobotov, V. A., "Artificial and Insights", AIAA Astrodynamics Oregon, 18-19 August, 1990
Section
Space Debris: Conference,
5.1
5.1
Monroe, Daryl; "ASE University of Texas
5.2
Philip G. Hill & Carl R. Peterson, Thermodynamics of Propulsion, 1965; pg. 371, 374, 490. Handbook
166M
5.3
CRC
5.4
Dr.
5.5
Daryl
5.6
"Evaluation of Advanced Advanced Space Analysis 12-13, 1988.
Westkaemper, Monroe;
Class
of Chemistry "ASE Graduate
Notes";
Mechanics and Addison-Wesley
and
376K
ASE-EM
Physics;
Class
Student;
Reference
University
Propulsion/ Power Office SVERDRUP/
99
60th
Department,
Publications;
Edition,
B-388
Notes" of
Texas
Concepts", NASA-LERC;
at
Austin.
April
5.7
David Kosmeyer; Graduate Student (University of Texas at Austin); Dissertation on Low-Thrust Electric Propulsion Option and Atmospheric Drag Effects.
5.8 "Electric and 5.9
5.10
Pradosh,
K.; vol.25,
William
"NASA
D.
1989,
pg.
"Characterization
Spacecraft,
Tushegee no.6,
Deininger
of Advanced Institute,
Nov.-Dec. & Robert
Electric
Propulsion
of
Propulsion
40-46. Electric
Alabama;
Journal
1988. Vondra;
"Arcjet
for SP-100 Flight Experiment; no.6, Nov.-Dec. 1988. Program";
Journal
AIAA
Propulsion Spacecraft;
Paper
87-1098,
1987.
"Performance July
5.13
Ray
Systems";
May 5.12
July-August
Propulsion
System vol.25, 5.11
Propulsion for Orbit Transfer"; Journal
Power;
of
10 KW
Xenon
Thruster";
NASA
TM
88-2192,
1988.
Bate, Roger Fundamentals
R.,
Donald D. Mueller, of Astrodynamics;
1971.
100
and Jerry E. White, Dover Publications,
New
York,
Section 5.14
5.2
Kohout, NASA p.
Lisa Lewis
L. and Faymon, Karl A. Space Power Research Center: Cleveland, February
14.
5.15
Ibid,
5.16
Faymon, Karl A. and Kohout, Technology for the Manned
p. 3.
and Energy Cleveland, 5.17
Storage January
Lisa L. Space Mars Mission
Systems). 22, 1986,
Lewis p. 27.
Power Systems (Pt. I-Photovoltaics
Research
Center:
Ibid.
5.18
Kohout, Lisa L. and Faymon, Karl A. Space Power Progress and Perspectives. NASA-Lewis Research Cleveland, April 4, 1988, pp. 2-3.
5.19
Baldwin, Richard S. (ed.). Space EleCtr0chemical Technology (SERT) 1989. NASA Lewis Research Cleveland, April 13, 1989, pp. 61-66.
Section
Technology Center:
Research Center:
and
5.4
5.20
National Facilities,
5.21
User's Guide for the Orbital Marshall Space Flight Center,
Section
Systems. 9, 1987,
Space NASA;
Transportation June, 1988;
System, p. 559.
Reference;
Maneuvering Alabama;
System
and
Vehicle; NASA June, 1989; p. 21.
5.6
5.22
"Space
Surveillance";
5.23
"IR Sensing Development;
5.24
Bachman, C. G., Laser Radar Systems House, Inc.: Dedham, MA, 1979.
5.25
Manhart, S. and Rangefinder for
will
Sky
be Tested May 1981.
& Telescope; on
July
Shuttle";
Industrial
and
P. Dyma, Self Calibrtating Space Application, Laser
101
1988. Research
Techniques.
&
Artech
Low-Power Laser Radar Technology
and
Applications. ed. by James Harney, TISOE: Bellingham,
M. Cruickshank WA, 1986.
and
Robert
C.
5.26
Bowman, S. R., Y. H. Shih, and C. O. Alley, Use of Geiger Mode Avalanche Photodiodes for Precise Laser Ranging at very low light levels, an experiment evaluation,. .....
5.27
Shapiro, Analysis
5.28
Erwin, H. O., "Laser Experiments". ....
Section
J. H., Robert W. for Peak-Detecting
Reinhold Laser
Docking
and D. Park, Radars"....
System
Radar
Wertz, Kluwer
James, R. Academic
5.30
Shank,
D.
and
Spacecraft Publishers:
Waligora,
Using Tracking and Data AAS/AIAA Astrodynamics Vail,
Colorado,
August
Attitude Boston,
S.
Relay Satellite Specialist 12-15,
5.32
National Facilities,
5.33
User's Guide for the Orbital Maneuvering Marshall Space Flight Center, Alabama;
5.35
H. Modern Sons: New
Spacecraft York, 1976.
Space Transportation NASA; June, 1988.
Orbit
and
Control.
Determination
(TDRS) Data". Conference,
1985.
Kaplan, Marshall John Wiley and
Section
Determination 1978.
"Automated
5.31
5.34
Flight
5.7
5.29
Section
"Performance
Dynamics
System,
Reference;
and
Control.
System
and
Vehicle; NASA June, 1989.
5.9 Crews, 1990.
Jeanne
Lee,
Personal
Communication,
November
19,
5.10 Hill, Philip G. & Peterson, Carl Thermodynamics of Propulsion, 1965; p. 371.
102
R.,
Mechanics and Addison-Wesley
Publications;
Section 6.1
6.3 User's Guide for the Orbital Maneuvering Marshall Space Flight Center, Alabama;
Section 7.1
7.0 "Final Design for a Comprehensive Program", STRES,Inc., University 1990;
7.2
7.3
pp.
Orbital of Texas
Debris Management at Austin; May 4,
60-72
"Techniques for Debris Control", Paper 90-1364, Petro, NASA Johnson Space Center, Houston, TX; AIAA/NASA/DOD Conference on Orbital Debris, 1990;
Baltimore,
Orbital
Debris
AIAA D. C., 7.4
Vehicle; NASA June, 1989; p. 26-30.
Progress 1989; pp
Andrew April
J. 16-19,
Maryland from
Upper-Stage
in Astronautics 201-213
NASA Engineering Exposition, Houston, Texas, October 29
103
Breakup, and
Johnson November
Joseph
Aeronautics,
Space Flight 1, 1990
P. Loftus, Washington
Center;
Jr.;
Appendix PFNO: DATE:
A00920 04/15/81
SUBSYSTEM TIER LEVEL TIER LEVEL TIER LEVEL
1 2 3
Anomaly
Spacecraft
SPACECRAFT: FLIGHT:
Reports LAUNCH: STATUS:
ISEE 3
08/12/78 UP
: INST-WIDENBCK : PRESSURE VESSEL : :
MISSION IMPACT POSSIBLE CAUSES ENVIRONMENT CODE OCCURENCE DURATION:
RATE:
IMMEDIATE LONG-TERM
RESPONSE SOLUTION
POSSIBLE
A
: : :
-
2 E M 4 4 : :
POTENTIAL
-
SLOW TOTAL
FOR
MAJOR
DEGRADATION LOSS (NO
IMPROVEMENT
D m
CAUSES: OPERATING TIME HUMAN/OPERATOR ERROR PROCEDURAL DESIGN OTHER UNKNOWN UNDEFINED
HARDWARE DESIGN MANUFACTURING WORKMANSHIP PART FAILURE MATERIALS INDUCED FAILURE ENVIRONMENTAL
J SYMPTOM
:
LEAK IN BER OVER TRAJECTORY COMMENT:
CAUSE
:
NOT KNOWN FOR SURE--PROBABLY DUE TO MICROMETEOROIDS SUFFICIENT SIZE & VELOCITY TO PUNCTURE THE 0.13MM BERYLLIUM-COPPER PRESSURE VESSEL WINDOW.
RECOVERY:
NONE
CORR.ACT:
USE
GENERAL OUR
GAS SYSTEM CAUSED COMPLETE A PERIOD OF ONE HALF HOUR. MEASUREMENT CAPABILITY.
LOSS OF GAS IN DRIFT CHAMTHIS CAUSES LOSS OF
OF
POSSIBLE. DIFF.DESIGN:NO-GAS
SYSTEMS
:
NOTE:
104
OR
BETTER
SHIELDING
OF
GAS
TNK
PFNO: DATE:
A00682 08/01/78
SUBSYSTEM TIER LEVEL TIER LEVEL TIER LEVEL
SPACECRAFT: FLIGHT:
: : : :
1 2 3
INST-HVESTADT PROPORTIONL LO-ENERGY
MISSION IMPACT POSSIBLE CAUSES ENVIRONMENT CODE OCCURENCE DURATION:
RATE:
IMMEDIATE LONG-TERM
RESPONSE SOLUTION
POSSIBLE
: 2 : E : MB
-
2 4 : :
ISEE 1
LAUNCH: STATUS:
CNTR DETCTR
POTENTIAL
-
FOR
INTERMITTENT TOTAL LOSS
MAJOR
(NO
IMPROVEMENT
D
CAUSES: HARDWARE DESIGN MANUFACTURING WORKMANSHIP PART FAILURE MATERIALS INDUCED FAILURE ENVIRONMENTAL
J SYMPTOM
:
SUDDEN COMMENT:
CAUSE
:
PROBABLY DUE MICRO-METEORITE.
RECOVERY:
NONE
LOSS
OPERATING TIME HUMAN/OPERATOR ERROR PROCEDURAL DESIGN OTHER UNKNOWN UNDEFINED
OF TO
GAS
PRESSURE
PUNCTURING
GENERAL
IN OF
POSSIBLE.
CORR.ACT:
OUR
10/22/77 UD
:
NOTE:
105
ONE THIN
OF
3
LOW
WINDOW(FRONT)
ENERGY
DETECTORS. BY
PFNO: DATE:
A00932 04/09/85
SUBSYSTEM TIER LEVEL TIER LEVEL TIER LEVEL
1 2 3
SPACECRAFT: FLIGHT:
LAUNCH: STATUS:
OCCURENCE DURATION:
RATE:
IMMEDIATE LONG-TERM
RESPONSE SOLUTION
: : :
-
2 D L
: :
POTENTIAL
5
-
C
-
FOR
MAJOR
SYSTEMATIC
CAUSES:
./
OPERATING TIME HUMAN/OPERATOR ERROR PROCEDURAL DESIGN OTHER UNKNOWN UNDEFINED
HARDWARE DESIGN MANUFACTURING WORKMANSHIP PART FAILURE MATERIALS INDUCED FAILURE ENVIRONMENTAL
SYMPTOM
:
CONTAMINATES SWITCH. THIS RESULTING IN COMMENT:
CAUSE
."
CONTAMINATES(PARTICLES} OF SWITCH MAY CAUSE OUTPUT.) RESTRICTED
RECOVERY:
ARE SUSPECTED TO BE WITHIN CLOSE CONDITION MAY CAUSE THE SWITCH LOSS OF KSA2 SERVICES.
OPERATION
IN VICINITY PARTICLES TO OF
WAVEGUIDE
CORR.ACT: GENERAL OUR
04/04/83 UD
: TLM & DH : LCP/RCP SWITCH : SA2 ANTENNA COMP :
MISSION IMPACT POSSIBLE CAUSES ENVIRONMENT CODE
POSSIBLE
TDRS 1
:
NOTE:
106
OF SWITCH. MIGRATE & SWITCH.
PROXIMITY TO BECOME
TO STUCK,
(CONTINUED DECREASE KSA
USE
PFNO: DATE:
0011 10/15/78
SUBSYSTEM TIER LEVEL TIER LEVEL TIER LEVEL
SPACECRAFT: FLIGHT:
TIROS N
LAUNCH: STATUS:
: THERMAL : : * : *
1 2 3
MISSION IMPACT POSSIBLE CAUSES ENVIRONMENT CODE OCCURENCE DURATION:
RATE:
IMMEDIATE LONG-TERM
RESPONSE SOLUTION
: : :
1 E L
-
MINOR
OR
NONE
m m
POSSIBLE
: D : *
-
CAUSES: HARDWARE DESIGN MANUFACTURING WORKMANSHIP PART FAILURE MATERIALS INDUCED FAILURE ENVIRONMENTAL
J SYMPTOM
:
CAUSE
:
RECOVERY:
THE TEMPERATURE PREDICTED.
THE WARMER CAUSED BY
OPERATING TIME HUMAN/OPERATOR ERROR PROCEDURAL DESIGN OTHER UNKNOWN UNDEFINED
OF
TEMPERATURE CONTAMINATION
THE
HYDRAZINE
GENERAL
COMPONENTS
OF THE HYDRAZINE OF THE THERMAL
CORR.ACT:
OUR
10/13/78 UD
:
NOTE:
107
IS
WARMER
TANKS AND COATINGS.
LINES
THAN
IS
PFNO: DATE:
41013 08/21/77
SUBSYSTEM TIER LEVEL TIER LEVEL TIER LEVEL
1 2 3
SPACECRAFT: FLIGHT:
: ARTICULATION : * : * : *
MISSION IMPACT POSSIBLE CAUSES ENVIRONMENT CODE OCCURENCE DURATION:
RATE:
IMMEDIATE LONG-TERM
RESPONSE SOLUTION
POSSIBLE
-
: 2 : S : M
VOYAGER 2
&
CONTROL
POTENTIAL
2
-
: A : A
-
CAUSE
08/20/77 UD
SUBSYSTEM
FOR
MAJOR
INTERMITTENT
CAUSES: OPERATING TIME HUMAN/OPERATOR ERROR PROCEDURAL DESIGN OTHER UNKNOWN UNDEFINED
HARDWARE DESIGN MANUFACTURING WORKMANSHIP PART FAILURE MATERIALS INDUCED FAILURE ENVIRONMENTAL
SYMPTOM
LAUNCH: STATUS:
:
:
PARTICLES EXCURSIONS CONTROL NIGHT'!
IN THE FIELD IN THE PITCH IN THOSE AXES.
OF VIEW OF THE TRACKER CAUSED LARGE AND YAW AXES WHILE UNDER CELESTIAL PROBLEM REFERRED TO AS THE 'BUMP IN
THE
WHEN THE CR240 ROUTINE IS EXECUTED THE PITCH AND YAW S.S. BIASES ARE INCREMENTED. THIS CAN HAPPEN EVERY .24 SEC. THEIR REMOVAL IS ONLY EVERY 1.2 SECONDS SO LARGE ERROR BIASES CAN ACCUMULATE. THE BRIGHT PARTICULE S CAUSE THE ERROR TO ACCUMULATE FASTER THAN EXPECTED. ALSO SEE PFR'S 3 7399,40411,40683
RECOVERY:
NONE
CORR.ACT:
THE SOFTWARE WAS REWRITTEN. INSTEAD OF INCREMENTING THE PITCH AND YAW BIASES TO FORCE A RELOAD OF THE S_S D_A CONVERTERS. THE CONVERTERS ARE . RELOADED DIRECTLY IN CR240 USING THE CORRECT VALUE (NO INCREMENT) OF THE S/S BIASES. THE CORRECTIVE S/W PATCH WAS LOADED TO BOTH FLIGHT S/C
GENERAL OUR
:
NOTE:
108
Appendix
B
B.1.1
SPinning
The
SPIDER
Maneuvering whether The
Other
design
should
SPIDER
satellites, targeted
region
SPIDER
will
will
rocket and
be
that
incorporates
made
vehicle
Since axis,
we
at the
the
same
three
robotic
either
place
as the
arms. a
itself.
thruster
firing
SPIDER
is reused, to
the
so
Space
it has
that
its
it will Station,
device
case,
the
orbit
will
be
able or
on
Remover
refueling.
109
about
to
it with
its
the
debris,
or
into
the
itself base,
SPIDER will
slowed
down
the
will
remain
atmosphere. from
for
major
spinning
on
be
a
by
the
will
large
debris
device,
detach
a similar
spinning the
the
The
DEbris
to
(inactive
collecting
the
decay to
for
OMV.
in
(m).
clamping
debris
the
debris
1 meter
itself
despun
for
atmosphere.
be
then
Orbital
considered
of
arms
probably
and
the
hardware)
robotic
attaching
thruster
In either
being
pieces
the
SPinning
debris
After
small
attached
return
SPIDER
large
than
will
to
attachment
into
three
B.1
the
rate
debris
greater
debris
foresee
track
with
large
it is
non-operational
the
Figure
similar
a modular
stages,
diameters
(SPIDER)
in fact
actively
send
Considered
a vehicle
and into
equipped
with
Options
Remover
(OMV), be
spent
debris,
DEbris
Vehicle
it
Design
debris
maintenance
by
a
If
the
and and
B.1.2
Tethering Tethering
is a concept that has been extensively
the last ten years [B.I:]. orbital
researched in
The principle of using tethers to eliminate
debris is to redistribute
the orbital
momentum of the debris.
Fuel is saved when the energy from the faster moving debris is used to increase the velocity for a propulsive
of the spacecraft, thus eliminating
the need
maneuver, while at the same time slowing the
debris down to a reentry orbit [B.2:].
Figure Calculations when
have
working km
altitude
for
each
deorbit
in
the
SPIDER
B.1.3
that
large
range),
design
in
earth
orbit
to
kilograms
low
eliminating The
been
Principle
a tether
debris
mission. has
Tethering
up
would
50
possibility
of using
be ( in
very the (kg)
efficient 200
km
of
fuel
a tethering
to
device
considered.
Netting Using
need
shown
with
700
B.2
for
tumbling
nets
despinning debris.
to
handle the
large
debris.
However,
and
small
It could
because
110
of
debris also
the
be
problem
would used of
eliminate to the
the
capture net
tearing
as well as potential danger to the spacecraft deploying
the net, we do
not believe it is feasible to net large objects.
Figure However,
this
objects, out
and
of
collect be
high
this any
created
to
SPECS,
current
of using
appears
with
the
the
SPIDER
best
idea
Design for
workable
fabrics
SPIDER
sized
Netting
believes
strength the
medium by
Inc.
be
B.3
has
debris
in
attaching
capturing
nets
could
like
Kevlar.
been
considered,
the to
area, the
medium be
Again,
and
fashioned the
possibility
principally any
sized
debris
to that
may
object.
References
B.1
Tethers in Advanced
B.2
Colombo, G., "The Use of Tethers for Payload NASA Contract NAS8-33691, Vol. II, March,
B.3
Carroll, Contract
Space Handbook. First Edition, Programs, January, 1985
J. A., "Guidebook for RH-394049, Martin
111
Analysis Marietta
NASA
Office
Orbital 19282
of Tether Corporation,
of
Transfer",
Applications", March,
1985
Appendix
C
Calculation
of
Perturbative
Thrust Acceleration Magnitude: • at = 4 Newtons / 21,000 Kilograms
lo
Solar
o
Pressure
Perturbative
Accelerations
= 2e-4
m/s 2
Acceleration
Magnitude:
• Compute Total Surface Area To Sun (Assume 50%) [ At] • Solar Arrays - Asa =2. (5" 14)= 140m 2 • Transfer Vehicle Body - Atvb =2. (.5"2)+(3" "Netting Modules Anm = 4 * ( 2 * 2.7 ) = 21.6 m 2
3)
= 11 m 2
• Propulsion Module Apm = 2.93 * 2 = 5.86 m 2 • Total Surface Area: At = Asa + Atvb + Anm + Apm = 178.46 m 2 • Total Area (cm) A = 1.7846e+6 cm. • Mass (gm) M = 2.1 e+7 gm • f =
-4.5e-5
Atmospheric • A/M = .085
a
• Compute • H = 400 • Space
• ra
=
• Va = 4.
J2
= 3.82414e-6
Drag
Perturbative
m/s 2
Station
=
Orbital
(-4.7034, 7.3853
5.488,
R V
a=6778.145 i=28.5, = =
(-4864.9, (-5.7619,
-1.5175)
km./sec.
Computed State = 2.5621e-5
Vectors: m/s 2
= =
1.
to Solar
Pressure:
2. Atmosphere 3. Oblateness:
Accelerations
Perturbative 5000
times
Drag: 4300 8 times
112
times
km.,
f_=300,
(-4864.9, (-5.7619,
Magnitude
kg/m
= 4.638e-8
Magnitude a=7540.645
R V
Vehicle. M=0
4555.153, -1281.306)km. 5.395, -1.5175) km./sec.
• adrag
Acceleration Orbital Elements:
• aJ2
Comparisons
m/s 2
km., e=0 _=0, w=0,
• p = le-12
i=29,
Thrust
= 3.8241e-8
Acceleration Magnitude: • Cd = 2.0
Elements:
State Vectors: 7.252e-5 rad/sec
Perturbative • Debris Torus •
• asp
state r and v vectors at Space Station orbit. kilometers is the lowest altitude for the Transfer
• Computed • Wearth
* A/M
3 (est.?) m/s 2
e=.l
w=200,
M=0
4555.153, -1281.306)km. 5.395, -1.5175) km./sec.
Appendix Satellite
D
Himawari
1
Rocket
Booster
Explosion
Data Type: Owner:
Delta US
Second
Launch
Date:
Stage
14.44
Jul
(2914) 1977
Dry Mass (kg): 900 (approx.) Main Body: Cylinder-Nozzle; 1.2 m by 5.8 m Major Appendages: Mini-skirt; 2.4 m by .3 m Attitude Control: None at time of the event Energy Event
Sources:
On-board
propellants,
Data Date: 14 Jul 1977 Time: 1612 GMT Altitude: 1450 km Location: Assessed
Post-Event
14N, 249E (dsc) Cause: Propulsion-related
Elements
Epoch: 77197.57445278 Right Ascension: 262.0317 Inclination: 29.0493 Eccentricity: .0973469 Arg. of Perigee: 66.7255 Mean Mean Mean Bstar: Cataloged
Anomaly: 303.2693 Motion Dot/2: .00007335 Motion Dot Dot/6: .0 .0 Debris
Debris Debris
Cataloged: in Orbit:
Maximum Maximum *Based
Cloud
delta delta on
Data 168 93 P: I:
937 3.0
uncataloged
min* deg* debris 113
data
range
safety
devices
Comments This
was
the
fragmentation.
It is
synchronous
orbit,
fragmented mission
fifth
on
also
the
day
had
Second only
energy
for
orbit.
The
the
40
of
propellants
burn.
The
(mainly elements
was
rocket
third the
stage
not
are
is
in a sun-
did
the
first
and
payload
after
available
which
perform
assessed
remaining the
a severe
burn,
body and
breakup
oxidizer)
above
experience
a depletion
This
the
to
which
performed
carrying
Earth
Stage
one
of launch.
low
depletion
the
which
successfully,
kg
Delta
its into
to have
after
Reference
Stage
J.R.; Rockets;
Headquarters
Explosion
of Satellite
Technical NORAD/ADCOM;
10704
Memorandum Colorado
114
and
81-5;
other DCS
Springs;
Delta Plans,
May,1981.
been
the
event.
Gabbard,
a
Second
the
Appendix
E
Fuel
Calculation
Program
Listing
PP,OGP, AM FUELC.OST REAL MNPIDRY,I'IPM.MNV.MA'.S'3.M20,M40,MgO,I_,p,p+IFUEL ,Pi2C,.N4Oq"+(.,, CHARACTER* I AN?w+_R I!10
PRINT
'+,'Inputnumb. +r of 20 cm
holes'
READ * ,h._') r-'PlP_/ _ ,h_i:,u_r,ur,,ber of 40 cn, holes' DEAD ++.N4O '"PINT * it+put number of qo on'.. P,ole5 _EAD * ,p+.o,+ pplUT _".:,++-,tltn',+LL_of Nett+r,9 P19tu_e (h', +:e! PEAC; '+.i'ir,IMDRY i-i
,I
+_
,-_,t, + *, Input rs:{55 Gt Pr,',_'.+;_i,,r, Module [ir, k
