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Dec 3, 1990 - Management Proposal. 8.1 . ...... system will then become the backup communications link between the TV and the ...... RECOVERY: CORR.

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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

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A

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N_A-_5382 zzr_ :1 l#1

Q

_NAL

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

,.- [email protected]

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