DECOM Mission

DEbris Capture and Orbital Manipulation

Department of Aeronautics & Astronautics

University of Washington

RASC-AL Team

Phillip Andrist, Alisha Babbitt, Vince Ethier, Michael Pfaff, Gabriella Rios-Georgio, T.R. Welter

DECOM Design Team

Romain Bertin, Sasan Boostani, Eric Braun, James Geier, Elizaveta Golaeva, Ben Grose,

Austin Lueck, Keith Neale, Nichole Sinarmanto, James Stuber, Jamie Waldock, Jasper Wang

Faculty Advisor

Arthur T. Mattick, Associate Professor

DECOM Mission

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Abstract Around 50 million objects, consisting of deactivated satellites, rocket bodies, and space system fragments orbit the

earth (1). The Kessler Syndrome predicts that as these objects collide and create more debris an exponential increase

in the total debris population occurs; eventually high traffic orbits such as LEO and GEO will be cluttered with

debris, rendering future space operations impractical.

DECOM Mission is designed to mitigate the Kessler Syndrome by actively removing five critical debris satellites

per year during a five year mission life. Residing within the 800-1010 km altitude and 81-83 deg inclination band,

the Cosmos Satellites, which use a Kaur-1 bus and weigh an average of 810 kg, are targeted due to their high

collisional probability. The DECOM Robotic Arm (DRA) captures, despins, and disposes of the Cosmos satellite.

Through mission heritage, the DRA can be modified to support removal of various debris types and future satellite

servicing. A NEXT Electric Propulsion engine, powered by UltraFlex solar arrays, performs debris transfer and orbit

lowering down to an altitude of 635 km, where the debris will be released assuring reentry within 25 years. The

mission takes advantage of differential precession between DECOM at the disposal altitude and the next debris

target altitude to achieve plane changes in right ascension, minimizing mission ∆V.

The estimated mission cost is $392M, a cost of $13M per debris removed. This mission approach is designed to be

sustainable for future debris removal missions; the DECOM Mission can actively remove 109 Cosmos, or 88 metric

tons of debris after four missions, with an average cost of $13M per debris object removed.

Introduction Space debris has been accumulating in Earth’s

orbit for around 55 years, and continues to grow

as yearly launches increase. Nearly 50,312,000

fragments of debris sized 0.1 cm and greater

orbit the Earth, totaling 1,900 tons of debris (2).

A small object with a 1 kg mass traveling at 10

km/s is capable of catastrophically breaking up a

1,000 kg satellite. Collisions such as these cause

more debris to be generated, leading to the

Kessler Syndrome, an exponential increase the

debris population in densely populated altitude

bands such as Low Earth Orbit (LEO) and

Geostationary Earth Orbit (GEO). It is estimated

that orbital debris costs $40 million per year in

mission failures due to irreparable satellite

damage, a number that will increase with the

debris population.

Many proposals for active orbital debris removal

missions have been advanced, but DECOM

Mission offers a 2015 launch-ready system that

is both cost effective and sustainable. The

DECOM Spacecraft is designed for close

approach rendezvous with a target debris object

and characterization of the dynamics of the

object using a long range camera, LIDAR, and

image recognition software. Capture is achieved

semi-autonomously with the DECOM Robotic

Arm (DRA) by a sequence of commands relayed

from Ground Control using the TDRS System

constellation. DECOM achieves control

authority over the coupled debris-satellite

system by determining the Cg via short pulses

from Attitude Determination and Control

System (ADCS) chemical thrusters and the

Inertial Measurement Unit (IMU). The NEXT

Engine then gimbals to fire through the new Cg

and performs an Edelbaum slow burn maneuver

to lower the coupled debris-satellite system.

Upon release of the debris at 635 km, an altitude

which ensures debris reentry within 25 years, the

satellite will then orbit raise to rendezvous with

the next debris target. During orbit lowering and

raising, the DECOM Spacecraft achieves the

right angle of the ascending node (Ω) of the next

debris target by utilizing differential precession.

This eliminates direct burns to accomplish plane

changes, significantly reducing required ∆V of

the total mission. The DECOM Mission profile

is shown in Appendix A.

The NEXT Engine has an Isp of 4100 seconds,

0.237 N thrust, and requires 6.9kW of power. A

total of 730 kg xenon is required to for a full five

year mission. The power system is designed to

generate a maximum of 14kW, by use of two

UltraFlex solar arrays with a total area 57 m2

and a Li-ion battery system capable of delivering

4.9kWh, 8kW power requirement during eclipse.

DECOM Mission

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The DECOM spacecraft, with an initial launch

mass of 2600kg, will be deployed with a Delta II

or Soyuz launch vehicle. A hexagonal isogrid

deck structure will support primary loads.

Through the implementation of high Technology

Readiness Level (TRL) components, an

effective solution to active debris removal was

achieved through the DECOM Mission.

Mission Definition The top level goal of the DECOM Mission is to

mitigate the Kessler Syndrome by stabilizing the

debris population with a mission that is cost-

effective, sustainable and launch-ready by 2015.

Previous studies have shown that removal of

five critical debris objects per year, starting in

the year 2020, will stabilize the debris

population (1). A key requirement of DECOM is

therefore to remove at least five objects per year.

Inclination Band and Altitude Range

The highest concentrations of critical debris

objects occur in the in the inclination ranges 71o

to 74o, 81

o to 83

o, and 96

o to100

o (3). The 81

o to

83o

inclination band was chosen for this mission

because approximately fifty percent of the high

mass objects are found in this region (1).

Confining the inclination region reduces the

mission ∆V by restricting inclination changes.

The two major altitude bands within this

inclination window that contain high

concentrations of critical debris are between 830

to 860 km, consisting of 67 objects totaling 240

metric tons, and between 950 - 1010 km,

consisting of 292 objects totaling 340 metric

tons (1). Lifetimes at this altitude can exceed

1000 years (4). Coupled with the high density of

active satellites in this LEO band, this region

contains the greatest potential to contribute to

the Kessler Syndrome. The DECOM Mission

will therefore remove debris from an altitude

between 800-1010 km and an inclination

between 81o - 83

o.

Debris Targets

Defunct satellites were targeted over rocket

bodies since developing the technologies to

rendezvous with tumbling satellites instead of

rocket bodies would be more applicable to for

future satellite servicing missions.

After eliminating rocket bodies, the list

consisted of Russian Meteor and Cosmos

satellites. The Meteor satellites were vastly

outnumbered by the Cosmos satellites and were

discounted to minimize ΔV and robotic arm

complexity. The main cluster of Meteor

satellites were located at an average inclination

of 81.2o compared to the average inclination of

83o of the Cosmos satellites. While 1.8

o plane

change seems small, removing the Meteor

satellites saved approximately 520 m/s of ΔV.

Furthermore, the Meteor satellites have two

solar panels external to the bus, which would

have greatly increased the complexity and

difficulty of capturing this object. Not targeting

the Meteor satellites enabled the arm length and

complexity to be reduced thus saving mass, cost,

and development time.

The vast majority of potential targets in this

inclination band and are Russian Cosmos

satellites and these will be the targets of the

DECOM Mission because of their large mass

and high probability of collision.

Cosmos is a general name for many different

types of Russian satellites such as Tsykada,

Tsiklon, Parus, Nadeshda and Zaliv satellites,

which include both military and civilian

satellites. These Cosmos use a KAUR-1

structural bus, which consists of a 2.035 m

diameter cylindrical spacecraft body and body

mounted solar cells and radiators. A 4.5 m long

stabilizing boom that is Nadir pointing stabilizes

the Cosmos spacecraft and reduces the

probability of three axis rotation. The KAUR-1

spacecraft body is shown in Figure 1.

Figure 1. A gravity gradient stabilized KAUR-1 satellite

DECOM Mission

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In order to target these Russian satellites,

international treaties and political negotiations

will need to be undertaken in parallel with the

construction of the spacecraft to ensure a 2015

launch date.

The Satellite Toolbox Kit (STK) Software was

used to identify an initial five year characteristic

mission list of 32 Cosmos satellites, which were

selected due to favorable distributions in Ω. The

order of the Cosmos satellites selected was

dictated by the relative change in Ω achieved by

the orbit lowering and raising maneuvers.

Target Dynamics and Charge Neutralization

Photometric flash data catalogues the rates of

light reflected off orbiting objects. This data

reports that approximately 73% of the Cosmos

targets exhibit only a steady reflection (5). This

indicates that these targets either have no spin,

or spin only about the nadir vector due to the

cylindrical design of the Kaur-1 bus. This data

further reports that Cosmos targets have a spin

decay time of 2 to 8 years due to natural

disturbance torques.

Surrounding plasma in LEO acts as a electric

ground for both DECOM and the targets. This

plasma charge is low in LEO except in polar

regions. As a result, DECOM will capture in low

latitudes during local daylight to decrease the

chances of charge buildup on either spacecraft.

A charge monitor device will verify this by

measuring DECOM’s relative potential to the

surrounding plasma just before capture.

Deorbit Altitude

The deorbit altitude was chosen to meet a 25-

year deorbit time to meet the post-mission

disposal requirements found in NASA-STD-

8719.1. Using the Debris Assessment Software®

(DAS 2.0), the parameters of the Cosmos

satellite, and the projected solar pressure flux for

2015 an altitude of 635 km was determined to

correspond to a 25-year reentry time.

Using STK it was determined that in 2015 there

will be 66 active satellites with altitudes less

than or equal to 600 km, and 596 known active

satellites between 600 km and 1050 km. There

will be 5653 total objects within the 600 to 1050

km altitude band compared to the 406 objects

below 600 km as seen in Figure 2. This

demonstrates that Cosmos satellites have a

decreased collision risk below a 600 km altitude.

Figure 2. Debris Population by Altitude

Debris Removal Trade Study There are many ways that debris can be removed

from a densely populated altitude. Moving the

debris into a designated graveyard orbit was

considered, but a major concern was that this

method was not sustainable. Over time, the

objects in this graveyard orbit will increase and

the overall debris problem will not be addressed.

In addition, the ∆V required to raise debris to a

2,000 km graveyard altitude requires more

than twice that to lower debris to 635 km. It was

therefore decided to lower the orbits of the

Cosmos satellites to an altitude where reentry

would occur within 25 years. Two

configurations were studied: a chemical or

tethered pod that would attach to the debris and

deorbit it and the use of the main spacecraft to

ferry the debris to a lower altitude and release it

using either electric propulsion (EP) or a

chemical thruster.

The pod configuration allowed the main

spacecraft to stay in the region of the Cosmos

satellites without performing significant orbit

lowering maneuvers. However, this

configuration is more complex, the pods are not

reusable, and a system to hold and attach the

pods is required. The main spacecraft orbit-

lowering configuration requires that the debris-

spacecraft system deorbit together, which

increases the time it takes to remove each debris

object. However, using an EP engine that both

transfers from object to object and transports the

debris to a lower altitude simplifies the overall

design.

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To quantitatively compare the configurations,

the pod system mass for a monopropellant

chemical thruster pod and a 3 km long tether pod

were calculated to compare to the propellant

required for the main spacecraft deorbit for both

a chemical and EP engine. The pod system mass

for the chemical thruster included the propellant

mass required to lower from 980 km to 600 km

and the engine mass. A chemical

monopropellant thruster and an EP engine were

used to calculate total system mass, including

propellant mass for orbit lowering from 980-600

km for a 16 object mission. The relative total

mass of both the propulsion system and

propellant mass per debris removed where

compared with respect to the main spacecraft

Electric Propulsion configuration as shown in

Table 1.

Table 1. Normalized System Mass per Debris Removed

Configuration Type Relative

Mass/

debris

Pod

Chemical Thruster 11

Terminator

Tether (3km) 2

Main

Spacecraft

Chemical

Thruster 14

Electric

Propulsion 1

The pod configuration adds significant system

complexity and would therefore only be selected

for this mission if it provided relative high

efficiencies in mass per debris object removed.

From the configuration trade study, it was

concluded that utilizing the main spacecraft

configuration with EP was the optimal solution

for lowering the debris orbit while using high

TRL technology available.

Transfer and Deorbit Maneuver The Cosmos satellites are distributed over the

360o range of Ω. Utilizing the main spacecraft as

a debris tug increases the ΔV requirements due

to the orbit lowering and raising for deorbit.

Instead of using direct engine burns to achieve

these new orbital elements, natural precession is

used. The Earth’s oblate shape causes slow

changes in the orbital elements of satellites in

time. The largest of these changes is the

precession of Ω and is given by Equation (1).

4

2 7

3cos

2

eRJ i

a

(1)

where J2 is a gravity harmonic, Re is the Earth’s

mean radius, a is the orbit’s semi-major axis and

i is the orbit inclination.

By lowering the altitude of DECOM as well as

the debris object, a differential precession rate is

achieved by DECOM relative to the debris near

a 980 km altitude. Because a low-thrust, high-

Isp engine is used for this deorbit maneuver, a

considerable amount of time is required to

execute the burn from 980 km to the deorbit

altitude of 635 km and back. Taking the launch

mass and the typical debris mass, these two

burns take a total time of 52.5 days, during

which relative precession is occurring. The

required time for transfer is calculated using the

typical low-thrust Edelbaum analysis for zero-

eccentricity orbits. The procession achieved

during this time is found by integrating Eq. (1)

with a time-dependent expression for the semi-

major axis. For the 52.5 day transfer, this

corresponds to a ΔΩ of 3.5°. This is the

minimum ΔΩ achievable for this transfer. To

achieve a larger ΔΩ, DECOM must coast at the

deorbit altitude until a ΔΩ in excess of 3.5° is

achieved. The values of 3.5° and 52.5 days are

subject to variability as the mass of the craft and

specific debris altitudes are taken into account.

The Edelbaum equation is used to generate ΔV

numbers for this transfer architecture. Because

on average the inclination changes are low, the

ΔV per burn is 180 m/s, which is the

difference between circular orbit velocities at the

debris altitude and the deorbit, or coast altitude.

Table 2. Characteristic 2015 mission

summarizes the characteristics of a 5-year

DECOM Mission assumed to commence in

2015, and for a launch mass of 2600kg. All

orbital perturbations were assumed to be

negligible except the effect of the Earth’s

oblateness.

DECOM Mission

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Table 2. Characteristic 2015 mission

Objects

Captured

Average

Time/Transfer

Total EP

Propellant

32 60.5 days 730 kg

These results demonstrate that DECOM is

capable of meeting the 5 object/year requirement

with the above transfer architecture. The

characteristic target list associated with the first

year of the 2015 mission is shown in Appendix

A. The table shows the targets located in the Ω

range in which DECOM will operate over the

first year.

In future missions operating in the same

inclination band, or in less populated inclination

bands, the average ΔΩ separation between

targets will be larger. If the time required in the

coast phase due to this increase in ΔΩ becomes

too large, the coast altitude can be lowered to

achieve a greater relative precession rate. This

decreases the time required for transfers at the

expense of additional ΔV and thus propellant.

The coast altitude, of course, could not be

lowered indefinitely. It has been determined that

the drag/thrust 1% at 400 km. Below this

altitude a drag model would need to be

implemented into mission calculations.

Rendezvous and Capture Capture of the Cosmos satellite will be

performed with the DECOM Robotic Arm

(DRA). The robotic arm was chosen over other

options because of its ability to capture non-

cooperative objects, TRL level, adaptability to

satellite servicing, moderate development cost,

and heritage.

DECOM Robotic Arm

The DRA, shown in Figure 3, is similar in

design to the European Robotic Arm (ERA) (6),

with roll, yaw, and pitch joints at the base, a

pitch elbow joint, and pitch and yaw joints at the

wrist. Using a design derived from ERA will

reduce development costs and increases the TRL

of the capture system.

Figure 3. DECOM Robotic Arm (DRA)

The DRA has a SARAH end-effector robotic

hand (7) and 6 degrees of freedom at 3 joints,

making the system dexterous and flexible

enough to capture various types of debris. The

arm has a maximum maneuvering extension of

3.5 meters and is a total of 4 meters long to

maintain a safe distance from the 4.5 meter long

boom on the Cosmos satellite. The arm is

projected to weigh approximately 315 kg,

primarily constructed of carbon composite

material. It will operate at a temperature range of

-35 to 75 degrees Celsius and with an average

operating power of 475 W and a thermal

hibernation power of 69 W.

Rendezvous with Debris Target

During rendezvous there will be three phases:

long approach, close approach, and capture

shown in Figure 4. Long approach will occur

after DECOM has finished its orbital transfer

maneuver and has arrived between 90m away

from the target. At this distance the long-range

LIDAR sensor and long-range camera will scan

for the target, record its rotation, and relay the

targets position and video to ground control,

where the dynamics of the Cosmos target will be

quantified.

DECOM Mission

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Ground control will determine if the spin rate

and geometry are within the capture capabilities

of the DECOM spacecraft.

Once confirmation is received, DECOM will

autonomously lower 10 m below the debris orbit

and coast to catch up to the target. When the

LIDAR sensor indicates that the DECOM

Spacecraft is 13 m from the target, the close

approach phase begins and DECOM will

autonomously reinsert itself into the debris’

orbit, 8 m from the target. At this point ground

crew will relay visual information on the chosen

capture point to DECOM’s onboard image

recognition software, along with secondary level

requirements for arm positioning during capture

to minimize torque loads and ensure safety of

the spacecraft. However, the final maneuvers of

DRA will be determined by the onboard

computer, taking into account the input from

ground. Capture will commence at low latitudes

on the daylight side of earth to prevent

hazardous charge transfer. The characteristics of

the debris target will be verified at this point;

updated DECOM Spacecraft positioning

commands and DRA maneuvering commands

will be relayed from Ground Control.

Having received the commands, the DECOM

Spacecraft autonomously thrusts toward the

target with a partially outstretched arm, retro-

fires to slow the approach, fully extends the

robotic arm, and captures the target with the

end-effector. Possible capture points are located

at the base, on the boom truss, or at the boom tip

as determined by the ground analysis of the

observation video.

The robotic arm will gradually stiffen and bring

DECOM and the target into co-rotation. Finally,

the ADCS system will perform a series of test-

firings to determine the new Cg of the system

and then despin and reorient the system to

perform the deorbit maneuvers.

Transient Arm Analysis

The coupled system dynamics immediately after

capture are complex and depend on how

DECOM chooses to capture the target. This

would likely not be determined until the spin

rates and orientation of a specific target are well

understood. Therefore, several simplifying

assumptions were necessary to estimate forces

and torques that would be transmitted through

the arm. These assumptions include:

1. Nadir-pointing Cosmos

2. 1 rpm pure nadir-spin only

3. Planar arm motion (3 DoF)

4. Boom truss capture point

5. Circular capture path

6. Azimuthal constant force at end-effector

Figure 4. Capture Procedure

DECOM Mission

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7. Capture in 90° rotation of Cosmos

8. Stationary DECOM

9. ICosmos = 830 kg m2

The analysis tracks forces and torques as a

function of the angular positions of the three arm

members. Angles were chosen which would

efficiently transfer the load at each point along

the capture path, constraining each joint to a

maximum rotation rate of 1.5°/sec as specified

by limits of the ERA joints. From this analysis,

the maximum forces and torques were calculated

at each of the 3 joints and at the connection of

the arm to the craft. The results are shown in

Table 3 where the joints are numbered below in

the order of increasing distance from the

spacecraft. Joint 5 corresponds to the wrist joint.

Table 3. DECOM Robotic Arm joint forces and torques

Joint Max Force, N Max Torque, N m

Body 9.4 31.3

3 9.4 28.2

4 9.2 14.5

5 8.9 3.9

The computed forces and torques are small

compared to the capabilities of the ERA. The

maximum allowable torques of the ERA joints

are 180 Nm at max speed (1.5°/sec) and 550 Nm

at min speed (0.001°/sec). The maximum brake

torque of 750 Nm (6).

Assuming the scaling of the ERA arm results in

a joint torque proportional to the mass ratio of

the two arms, this results in maximum torques of

90 Nm, 225 Nm and 325 Nm for maximum

speed, minimum speed, and braking,

respectively. It should be noted that torques

applied to the body are not influenced by these

limits as this is not an arm joint. Thus, the arm

can generate greater than three times the torques

necessary to achieve this arm trajectory, and the

brakes would not slip unless the torques were

more than 10 times greater. The torques and

forces transmitted through the body are

significantly less than the loads during launch,

and thus are not drivers for structural design.

Taking these forces as typical for a capture

scenario in which DECOM is free to move,

these forces exert rotational and translational

accelerations on the craft. These are negligible

relative to launch accelerations, but unlike

launch, the solar panels are deployed. Structural

analysis of the deployed solar arrays during

capture accelerations show a frequency response

greater than 1Hz.

Propulsion System Due to the high ΔV requirements of this

mission, an electric propulsion (EP) engine was

chosen after analyzing a variety of appropriate

high Isp engines. The EP thruster selected was

the NASA Evolutionary Xenon Thruster

(NEXT) (8). At the thruster’s full power mode

of 6900 Watts, it produces 0.237 N of thrust at

4100 seconds specific impulse. This ion engine

was tested, developed and built by NASA and

Aerojet. The NEXT engine is TRL 6 due to

multiple engineering and lab models that have

been tested. Furthermore, its predecessor, the

NSTAR, is currently flying onboard the DAWN

spacecraft.

During the five year mission, the NEXT would

have a total throughput of approximately 730 kg

of xenon. Current models of the engine indicate

that the first mode of failure for the ion

accelerator grid occurs at 750 kg throughput of

xenon gas. Three tanks circumferentially

mounted in the center of the spacecraft will hold

810 kg of propellant, allowing for ullage and

specific impulse degradation over the life of the

thruster.

Attitude Determination and Control System

The Attitude Determination and Control System

(ADCS) was designed to maintain the DECOM

spacecraft at the required attitude for thrust and

solar panel orientation, despin of the coupled

debris target and DECOM system, and to

overcome disturbance torques that the spacecraft

would experience during its mission. This

requires the ability to provide fine control when

approaching the target as well as the ability to

exert a relatively high torque on the coupled

system.

For fine attitude control and counteracting small

disturbances, three orthogonal L-3

Communications MWA-50 momentum wheels

(9) are used for fine attitude control.

DECOM Mission

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The thrust requirements of ADCS are set by the

time allotted for the coupled despin maneuver.

Assuming an initial Cosmos rotation of 1 rpm

about three orthogonal axes, a co-axial capture,

and a distance from DECOM Cg to the target Cg

of 8m, the thrust required to accomplish

complete despin within 20 minutes was found to

be approximately 10N. To conduct regular

mission maneuvers and these despin operations,

16 Northrop Grumman MRE-4.0 hydrazine

thrusters are used. Each thruster has a specific

impulse of 217 seconds and a thrust of 9.8 N.

These thrusters are also used to desaturate the

momentum wheels. Shown in Figure 5. ADCS Thruster Location, four of these thrusters are

placed at each end of the spacecraft so as to

provide full control over pitch and yaw and four

thrusters are placed on the arm end and at the

NEXT end of the spacecraft, pointed

perpendicular to the support structure of the

solar panels. This allows movement in the

direction perpendicular to the thrust vector in

addition to roll control

Figure 5. ADCS Thruster Location

Propellant Budget

During the time spent precessing to the correct

right ascension, the worst case scenario would

have the ADCS performing up to 876 firings per

debris target. The hydrazine propellant budget

for the ADCS system is summarized in Table 4.

Table 4. ADCS Mass Budget

Maneuver Thrust

Propellant

Mass/

debris

Transfer Momentum wheel

desaturation 2.5 kg

Capture

and

Rendezvous

Approach 0.2 kg

Close Approach 2.6 kg

Cg Determination 0.1 kg

Despin 0.2 kg

Reorientation 0.2 kg

Maneuvering Sub-total 5.8 kg

Total w/Margin 7 kg

Tank Capacity (12% Margin) 250 kg

An ATK 80514-1 bladder tank was chosen to

hold 250 kg of hydrazine, which is sufficient to

support ADCS control for 35 debris targets over

a five year mission. Since we will not be

targeting 35 objects the remaining balance of

this propellant can be used to balance the

spacecraft for launch. An ATK 80458-201 tank

is used to store the nitrogen gas used for

pressurizing the hydrazine. This nitrogen

pressurant is stored at 2,875 psig and is

regulated to the required hydrazine pressure of

275 psia.

Power System Power consumption of principal components of

DECOM is listed in Appendix C. The NEXT

thruster dominates consumption, and the orbit-

average power is 8 kW. This power will be

generated using two solar array wings in

sunlight, and by batteries during eclipse and

during rendezvous, when orientation of solar

panels might otherwise constrain capture

maneuvers. A simulation of the eclipse

environment showed that the maximum eclipse

time is 35 min, while the average orbit period is

95 min. Because of the high inclination most

eclipses will have duration of 29 min, and ~20%

of orbits will have no eclipse at all. To provide

full power during the longest eclipse, and keep

average depth of discharge at 27%, the storage

capacity of the battery system must be 4.9 kW-

hr. The batteries will be charged during sunlight

DECOM Mission

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(4.7 kW) from the solar array, requiring a

maximum array power to be 12.7 kW.

Figure 6. UltraFlex Solar panel

Because of their resistance to oscillations due to

spacecraft maneuvers when compared to

rectangular arrays, the UltraFlex solar panel

design (10) will be used, as depicted Figure 6.

This design uses a circular array of 6-m

diameter, which can produce 7 kW, or 14 kW

for two arrays, providing a 10% margin above

required power. This design is to be used on the

Mars Phoenix Lander and has a TRL level of 7

(11). The total array area is 57 m2 and mass is

90 kg. A main bus voltage of 100 V was chosen

to keep bus current moderate and connecting

wire less massive; the NEXT power control unit

is designed for a 100-V input. The arrays will

have a stowed position for launch that will

include a latch type device. The array support is

designed to accommodate the forces and torques

experienced during capture of a target spacecraft

with the arrays fully deployed. The panel

support will allow rotation about its axis as well

as the roll axis of the spacecraft to insure the

panels can maintain a sun-facing orientation.

The battery system will use Li-ion batteries; to

keep the number as small as possible, the

largest-capacity space-rated batteries from Saft

were chosen. The Model VES-180 cells, with

nominal voltage of 3.6 V, 50 A-hr charge

capacities have a mass of 1.1 kg. The battery

pack will use 3 series-connected strings of 28

cells to achieve the required bus voltage and

energy capacity. Six spare cells are included to

replace any malfunctioning battery, for a total of

90 cells and a battery system mass of ~100 kg.

Wiring and auxiliary power system components

were assumed to have a mass of 40% of the

arrays and batteries, bringing the total power

system mass to ~ 266 kg. The power

distribution system is illustrated in Appendix C.

Communications & Data Handling The data handling system of DECOM is

illustrated in Appendix B. This system receives

inputs from navigation and health-monitoring

sensors to establish the state of the spacecraft,

and implements controls software to activate

ADCS, power and propulsion systems to

maintain the desired orbit. This system also

relays information to ground control stations and

receives commands from ground during the

mission. For most of the mission, when

DECOM is acquiring the orbit of a debris target

or deorbiting a debris, the data rates are

estimated to be ≤ 1 kBit/sec. The most data-

intensive part of the mission is during

rendezvous and capture, when live video of the

debris is relayed to ground for assessment and

determination of propulsion and robotic arm

maneuvers needed to capture the target. The

data rate is estimated to be ≤ 3 MBit/sec during

this period, which may last up to 1 day.

To insure continuous communication during this

period, DECOM will utilize the TDRS system

for communication (12) illustrated in Figure 7.

Figure 7. TDRSS Communication Diagram (11)

DECOM Mission

10

This system has multiple-access (MA) channels

that can accommodate data rates up to 300

kBit/sec, and single-access (SA) channels for

rates up to 6 Mbit/sec on S-band. MA channels

(2.1046 GHz forward service, 2.2875 GHz

return) can be used for the low-bit-rate part of

the mission, and it is planned to use the S-band

SA service (2.075 GHz forward, 2.25 GHz

return) during rendezvous. From the TDRSS

manual and assuming 3dB losses due to pointing

inaccuracy and polarization losses, the required

EIRP for MA at 1 kBit/sec is EIRPMA=6.4dBW

(4.4W), and for SA at 3 Mbit/sec,

EIRPSA=27.4dBW (550W) (12). A low-gain

omnidirectional antenna (patch) will suffice for

MA service, and it is planned to use patch

antennas on each of the 6 side faces of DECOM

for redundancy. For the SA service, a medium

gain antenna (gain ~20 dBi) is required to keep

the communication power reasonable. To avoid

the need for mechanically pointing a parabolic

antenna, it is planned to use phased array

antennas on DECOM, which can be

electronically steered. The antenna design is a

7x7 array of patch elements, spaced by ½

wavelength (7 cm), resulting in a planform of ~

0.5 m x 0.5m. Phase variation to achieve

desired pointing is accomplished by 4-bit phase

shifters. A simulation has shown that the peak

gain is 21.2 dBi at broadside, falling to 18 dBi at

a pointing angle of 60°. Furthermore, it was

found that at an altitude of 980 km (the average

for Cosmos debris) where high-rate

communication is needed, two phased array

antennas mounted on faces adjacent to one of

the solar-panel faces can provide a nearly

continuous (99.7%) link assuming each antenna

can be directed within a 60° cone. The

transmitter output will be switched to the

antenna with the better link (smaller angle)

during SA communication with TDRSS. At the

steepest angle of 60° the minimum radiated

power must then exceed 9W; to allow margin

for unforeseen losses, the design radiated power

is 20W. Assuming dissipation and reflection

losses from the antenna of 60% the transmitter

output must be 50W, and assuming a transmitter

efficiency of 25%, the DC power input to the

transmitter will be ~200W.

The DECOM Data Management Computer

(DMC) will act as the interpreter and

communicator of data to and from the DECOM

Robotic Arm, Navigation equipment, pressure,

temperature, and systems status monitoring

equipment aboard DECOM in order to maintain

safe operation of the autonomous systems

aboard DECOM as well as transfer the

information to the Ground Station for analysis

and control. The DRA video feed will be

compressed by the DRA main processor and

sent to the DMC via an input bus that will also

receive inputs from the other DECOM systems

and will be required to process the telemetry

signal of up to 3 Mbit/sec. The DMC will select

the appropriate antenna based on mission mode

data rate, attitude and TDRSS in order to

transfer data at scheduled times to the TDRSS

Satellites via the transmitter amplifier and

antenna (13).

Navigation Equipment

The primary navigation equipment used during

transfer operations are a star tracker and a GPS

receiver. The Space Micro Inc. SM-MDE1300

star tracker was chosen for its simplicity,

compact size, low power draw, and attitude

determination accuracy ranging from 1 to 10

arc-seconds (14). The Honeywell Enhanced

Space Integrated GPS/IN (E-SIGI) provides

better than 50 m position accuracy and the

velocity reading is better than 0.3 m/s RMS

(15). To ensure the safety of the DECOM

mission during rendezvous TDRSS will be used

to obtain position accuracies to within 30 cm.

Earth sensors onboard provide redundancy for

spacecraft attitude determination and will be

used to reorient the spacecraft in anomalous

situations when orientation is unknown.

Structural Design & Analysis Launch Vehicle Selection of the Launch Vehicle (LV) upon

which DECOM would be sent to orbit was an

important design decision that needed to be

completed early in the design process, due to

many design characteristics being directly driven

by the LV specifications, interfacing, and launch

load environments. Key mission profile

parameters were used to narrow down the

possible LV systems.

DECOM Mission

11

DECOM will be operating at a very steep

inclination band of 82.9° and an initial altitude

of ~900km. In the interest of propellant savings,

a LV system which could place the spacecraft

here initially was desired. Using an initial launch

mass estimate between 2500-3000kg which was

estimated from relations to the payload mass

(16), a trade study was performed on the

available market LV systems which could carry

this mass to this orbit. As a result of this study,

the Delta II 7300 was chosen as the primary

design vehicle, with the Soyuz as the backup

system. The interface requirements for both

LV’s were taken into consideration for the

spacecraft design, in the chance that one LV

system would not be available to fly.

Launch environments aboard the Delta II are

typically larger than that of the Soyuz, and thus

were used as the limit load envelope for the

structural design. A summary of the expected

load environment through LV Ascent and

MECO is provided in Appendix D.

Primary Structure The driving configuration constraints for the

spacecraft were optimal usage of the payload

fairing volume, and the storage of the robotic

arm and the large UltraFlex solar arrays. The

dimensions of the UltraFlex solar panels in their

stored launch-configuration dictated the

minimum height of the spacecraft. Also, the

dimensions of the robotic arm members set

minimums for the spacecraft sizing in order to

provide safe storage. To accommodate these,

and maximize the usage of the PLF volume, a

hexagonal buss structure was chosen, with a 4m

height and 1m face lengths.

The primary structure of DECOM was designed

to transmit launch loads through the payload

adapter fairing (PAF), while supporting the most

massive and critical components aboard the

spacecraft. The primary axial loads are directed

through six reinforcing stringers oriented at the

corners of the hexagonal prism, which interfaces

with the LV PAF by a pin-joint truss structure.

Loads from internal components, such as the

propellant tanks and the robotic arm, are

transmitted to this structure through a similar

pinned truss structure, while secondary

components are secured to either iso-grid decks

or the spacecraft facing panels.

To evaluate the structure design, the Finite

element computational program ANSYS (17)

was utilized. A finite element analysis of this

structure under the maximum aerodynamic

loading during LV ascent (3.125g’s laterally and

3.25g’s axially) is shown in Figure 8.

Figure 8. FEA Analysis During Launch

Component Layout

Aluminum honeycomb panels are used along the

six sides of the spacecraft running parallel to its

axis. These panels serve to increase the bending

stiffness and stability of the spacecraft, as well

as to serve as emissive surfaces for thermal

control of the spacecraft interior. A thermal

analysis was performed using Thermica software

(18) and is presented in Appendix D.

Considerations for component placement were

primarily driven by sensor field of view

requirements, maintaining clearance from the

NEXT engine exhaust plume, and the robotic

arm arm’s operational region. Moreover, the

height of the center of mass above the LV PAF

separation plane had to meet torque

requirements.

DECOM Mission

12

Figure 9. DECOM Component Configuration

Figure 9 shows the final configuration of

DECOM’s components. The final DECOM

itemized component list and mass budget are

presented in Appendix D. The DRA in stowed

launch configuration is shown in Figure 10.

Figure 10. DRA in Stowed Configuration

Solar Array Structural Analysis

DECOM will be encountering many impulse

and shock loadings throughout the rendezvous,

attitude control, and release mission modes.

Traditional solar arrays are very sensitive to

vibration responses because they tend to have

poor dynamic damping and under some

circumstances cause permanent damage to the

array. Therefore, most systems avoid high

impulse loading when solar panels are fully

deployed. This, however, is not an option for the

DECOM mission due to the rendezvous phase.

The UltraFlex solar arrays have relatively high

fundamental mode frequencies and superior

damping characteristics compared to more

traditional solar panel designs. This is in part

due to their use of a Vectran open mesh

substrate, which reduces the amount of non-

power producing mass present in the array, and

thus less momentum stored within the panels, as

well as due to their radial truss structure.

An FEA was performed on the solar arrays,

which indicated their first fundamental mode to

occur at a frequency of 5.9 Hz with a deflection

of no more than 1.5 cm. This is far less than the

expected impulse vibrations DECOM is

expected to encounter. Forces during rendezvous

procedures result in a 5-10N-m of torque about

the solar panel joints. This force is negligible

relative to the material strength of the joints,

with a factor of safety of ~85.

Mission Effectiveness Sustainability

A mission analysis was performed to determine

how many times the DECOM mission could be

performed before the targets in the chosen

inclination and altitude became too sparse for

the current mission architecture to be feasible.

The analysis was carried out using STK to

generate a mission list for the years 2015, 2020,

2025, and 2030 and removing the captured

targets from previous years. Table 5 presents the

findings, which show that four missions are

viable in the 81o to 83

o and 800 to 1050 km

region, and gives the number of targets captured

per year, average time per object, and the xenon

propellant mass per mission. The times

presented only represent transfer time and not

rendezvous time, which is expected to be no

more than 1 day per object. Mission four would

not meet a strict five object per year

requirement. However, looking at the four

missions collectively, 109 objects,

corresponding to over 88 metric tons, would be

removed over 20 years. This corresponds to

approximately 5.5 objects per year, which shows

that over the duration of this mission, the

requirement is met.

DECOM Mission

13

Table 5. 20-year Mission Results

Mission

Year

Targets

Removed

Average

Time per

Object

(days)

Xenon

Mass

(kg)

2015 32 60.4 729

2020 30 63.4 704

2025 26 73.2 727

2030 21 91.6 715

In the 2020, 2025, and 2030 missions, a lower

coast altitude is used to maximize the number of

targets removed for a fixed propellant budget.

Other possible mission modifications which

would help accommodate a sparse debris

population include using multiple engines or an

engine with higher thrust. Future missions

would most likely incorporate technological

advancements in both thruster capabilities and

solar panel efficiencies. For wide enough

distributions of Ω, it will no longer be feasible to

meet the five objects removed per year with one

spacecraft. To overcome this, multiple

spacecraft could work in parallel to deorbit five

objects per year, if required. While there are

numerous Cosmos in orbit, a modified robotic

arm design would be needed to capture more

complex spacecraft and possibly rocket bodies.

These modifications have direct applications to

future satellite servicing missions.

Mission Life and Cost Budget

A cost budget performed using NASA and U.S.

Air Force cost estimating relationships (CERs)

(19), estimates the cost work breakdown

structure (WBS) of the DECOM Mission, shown

in Table 6. The DRA, UltraFlex solar arrays, and

NEXT Engine test units will be used as

protoflight qualification. New software will have

to be developed to support the image recognition

software and capture real-time sequencing.

Table 6. First Mission Budget Estimate

WBS Cost

(FY2015) $M

Space

Segment

RDT&E 92.2

Software 56.9

TFU 98.2

Launch

Segment Delta II LV 68.3

Ground

Segment

Ground Station 79.3

Operations and

Maintenance 52.4

Lifecycle 392

The DECOM Mission was designed to minimize

the cost per debris object removed. Shown in

Table 7 is the cost per debris over the four, 5-

year DECOM missions. Shown in Appendix A

is the cost trend over the 20 year total mission

life cycle.

Table 7. Mission Cost Characteristics

Total

Cost

(FY2015$M

)

Total

debris

removed

Total

Cost per

debris

removed

(FY2015$M)

1st

Mission 392 32 12.2

2nd

Mission 366 62 12.2

3rd

Mission 344 88 12.5

4th

Mission 326 109 13.1

TRL and Mission Life

The TRL of components was a large factor in

choosing certain designs to meet the launch-

ready year of 2015. The UltraFlex solar arrays,

DRA, and NEXT are technologies that all have

flight proven predecessors. While this means

that these technologies must be flight qualified,

it also demonstrates that the DECOM Mission

will be a development platform that will be a

cutting-edge mission by year 2015. The TRL of

the major components are shown in Table 8. The

DECOM Mission

14

highest risk technology is the NEXT Engine,

which reaches an extrapolated failure mode near

the end of the first mission.

Table 8. TRL of Critical Components

System Component TRL

Propulsion NEXT Engine 6

Propulsion Gimbal Mechanism 5

Power UltraFlex 7

Capture DRA 6

Robotic Arm Prototype

A robotic arm prototype was constructed to give

students experience with manufacturing, plant

modeling, and designing control systems. The

functionality of the capture portion of the

DECOM mission was emulated. In practice, a

user inputs the position of the object and the

desired position of release through the use of a

Graphical User Interface (GUI). This emulates

the DECOM rendezvous procedures where the

object recognition portion occurs on Earth and

the position of the hard point for capture is sent

to the robotic arm. The prototype is shown in

Figure 11.

Figure 11. Prototype Robotic Arm

Prototype Arm Control

The controls problem was to accurately position

the hand effecter to capture a model debris

object. The controller also needed to ensure the

safety of the robotic arm at all times as well as

move the captured debris object to a specified

release location. Solving this problem required

accurate modeling of the plant for simulation

testing. Both the prototype and DRA were

modeled in a Matlab toolbox called

SimMechanics. The mission controller handled

the inverse kinematics, conflict avoidance

algorithms, and path planning. A feedback

position controller was designed to achieve the

goal of capturing an object. Hardware-in-the-

loop testing for the prototype arm commenced

once satisfactory simulation results were

achieved.

Mission Modifications

The DRA is a six degree of freedom arm with

much greater complexity than the robotic arm

prototype which has four degrees of freedom.

The increased freedom increases the complexity

of the plant models as well as solving the inverse

kinematics equations. Movements of the arm

will result in torques on the spacecraft which

will have to be counteracted by the ADCS

system. The robotic arm prototype was not

designed to operate in a zero-gravity space

environment and therefore the actuators and

sensors used only mimic the functionality

required for the DRA. For example, IR

proximity sensors were used to detect relative

distance of obstacles as well as the distance to

the object being captured. On the DECOM arm,

LIDAR will be used to detect the relative

distances for object avoidance and a camera

would be used to observe the hard point being

captured.

Outreach The outreach program consisted of DECOM

members volunteering and presenting at the

Seattle Science Center for Polar Science

Weekend, the University of Washington for the

National Amateur Rocketeers Conference

(NARCON), and at Seattle Central Community

College. These events generally consisted of

science demonstrations and conversations with

children and their parents about STEM topics. In

addition, the DECOM Mission and the orbital

debris problem were presented at the University

of Washington for Engineering Discovery Days,

where an estimated 20,000 prospective students

and students from local grade schools attended

this year. The main focus of the DECOM

outreach program was on building and teaching

a curriculum that introduced space, basic

engineering design and simplified rocket

DECOM Mission

15

dynamics to two classrooms of students at

Maple Elementary School. The full curriculum

and worksheets are presented in Appendix H.

During the first two visits, DECOM members

presented topics and discussed them with the

class in interactive workshops. The third visit

involved launching water bottle rockets and

exploring how to design for distance versus

height while briefly explaining momentum

transfer in simplified demonstrations.

Figure 12. Maple Elementary Outreach

A blog about the UW Senior Space Design

DECOM mission was maintained for the

purpose of bringing interested students along our

design process and demonstrating challenges

and solutions. (20). The DECOM team

contributed approximately 180 hours and

reached about 1300 people.

Conclusion The DECOM Mission provides a cost-effective,

sustainable method of actively removing critical

space debris that can be operational by the year

2015. DECOM performs a coupled transfer and

deorbit maneuver to reduce mission ∆V by using

a NEXT Electric Propulsion thruster, powered

by two UltraFlex solar arrays. The DECOM

Robotic Arm (DRA) and rendezvous system

designed to capture the debris will develop the

technology for future satellite servicing and

active debris removal systems. The DECOM

spacecraft is capable of removing 32 debris

targets within the first 5 year mission and can be

operated without system alterations for four

missions, resulting in a total of 109 objects, or

88 metric tons of debris removed. The first

mission cost is approximately $392M and the

total lifecycle cost for four missions is $1.5B,

resulting in an average of $13M per debris

object removed. The DECOM Mission

combines moderate-to-high level TRL

technologies that are deployable by 2015 and

will be evolved over the mission life, providing

both a platform to develop new technologies and

to stabilize the orbital debris population in Low

Earth Orbit.

Bibliography 1. The top 10 questions for active debris removal. Liou, J.-C. Johnson : NASA, 2010. European

Workshop on Active Debris Removal.

2. Space Surveillance/Space Debris. Australian Space Academy. [Online] 05 18, 2010.

http://www.spaceacademy.net.au/index.htm.

3. Carrol, Joseph A. Space Transport Development Using Orbital Debris. Final Report on NIAC Phase

I. [Online] [Cited: January 22, 2011.] http://www.spaceelevator.com/docs/800Carroll.pdf.

4. Wertz, James R. and Larson, Wiley J. Sapce Mission Analysis and Design. Springer : Microcosm,

1999.

5. McCants, Mike. PPAS (Photometric Periods of Artificial Satellites). [Online] 3 2011.

http://www.satobs.org/tumble/tumbleintro.html.

6. European Space Agency. European Robotic Arm. Leiden, Netherlands : ERASMUS User Center and

Communication Office, 2007.

7. Novel Robotic Hand SARAH For Operations on International Space Station. Rubringer, Bruno. 2002.

8. NEXT Ion Propulsion System Development Status and Capabilities. Benson, Michael J. Patterson

and Scott W. Cleveland, Ohio : NASA Glenn Research Center, 2003.

9. Space and Navigation . MWA-50 Low Cost Momentum Wheel Assembly. Budd Lake, NJ : L3, 2008.

10. Brian S. Smith, P Alan Jones, Stephen F White, T Jeffery Harvey. High Speciic Power Solar

Array for Low to Mid Power Spacecraft. Sylmar, California : Spectrolab, 2007.

11. 19th Space Photovoltaic and Technology Conference. NASA Glenn Research Center. Brook Park,

Ohio : NASA, 2007. NASA/CP-2007-214494.

12. NASA Goddard Spaceflight Center. Space Network Users' Guide. Greenbelt, Maryland : GSFC,

2007. 450-SNUG.

13. Space Science and Engineering. Space Qualified Single Board Computers. 2007. SPARC V7.

14. INC., Space Micro. Space Radiation Hardened Star Tracker. SpaceMicro.com. [Online] [Cited: 2 22,

2011.] http://www.spacemicro.com/pdfs/star_tracker_v2_7.pdf.

15. Honeywell. E-SIGI Enhanced Space Integrated GPS/INS. [Online] [Cited: 2 22, 2011.]

http://www51.honeywell.com/aero/common/documents/myaerospacecatalog-documents/E-SIGI.pdf.

16. Larson, James R. Wertz and Wiley J. Space Mission Analysis and Design. s.l. : Microcosm Press,

1999.

17. ANSYS . ANSYS 13.0 . Canonsburg, PA : s.n., 2010.

DECOM Mission

17

18. MSC Software . THERMICA: An Integrated Thermal Design Environment. Santa Ana. CA : MSC

Software Corporation, 2008.

19. Unmanned Space Vehicle Cost Model. s.l. : SMC, 2000.

20. DECOM Mission. UW Orbital Debris Mitigation. [Online] University of Washington, 2011.

https://sites.google.com/site/sendebris/.

21. Verhoeven, Didier. Space Mechanisms and Tribology for the Joints of the European Robotic Arm.

Belgium : s.n., 1999.

DECOM Mission

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Appendix A Mission Summary

The DECOM Spacecraft will launch via a Delta II or Soyuz launch vehicle. DECOM will then

separate from the payload fairing, deploy the UltraFlex solar arrays and begin entering into

transit for rendezvous with the first Cosmos target. The DECOM Robotic Arm will extend and

capture the target at an altitude of approximately 950 km. Upon capture, the DECOM ADCS

system will despin and reorient the coupled DECOM-target system. An Edelbaun maneuver will

be performed with the NEXT Ion Engine to orbit lower to 635 km. The target will be released

and DECOM will reorient to rendezvous to the next target.

Figure 13. DECOM Mission Profile

Table 9 the target list associated with the first year of the 2015 mission. The mission is calculated

only taking into account relative precession between DECOM and targets in the inclination band,

but neglects the rate at which these targets precess relative to one another.

Captured targets are in gray, while white targets must be skipped because the cumulative ∆Ω is

smaller than the ∆Ω achieved by transferring to the deorbit altitude alone. The cumulative ∆Ω is

the ∆Ω between the current target and the previous captured target.

DECOM Mission

19

Table 9. Characteristic mission: first year target list. Captured targets in gray.

Mass, kg Altitude, km Eccentricity Inclination Ω Cumulative

ΔΩ

825 969 2.9E-03 82.91 336.09

810 975 3.6E-03 82.90 333.59 -2.50

810 978 5.0E-03 82.88 333.13 -2.96

810 973 2.1E-03 82.96 331.71 -4.38

810 977 3.4E-03 82.89 325.82 -5.89

825 988 3.5E-03 82.89 323.40 -2.41

810 975 6.0E-03 82.88 318.10 -7.72

810 981 4.2E-03 82.90 318.06 -0.05

810 989 3.5E-03 82.89 317.11 -0.99

825 978 3.5E-03 82.87 315.23 -2.87

810 964 1.6E-03 82.88 314.21 -3.89

A parametric cost model was used to estimate the mission cost and cost per debris object

removed for all four missions. Figure 14 shows the total cost of all four missions by using the

estimate that 60% of the cost will be spent during the first 2.5 years of the mission. The cost per

deorbit corresponds to the projected number of targets DECOM mission can remove for each

mission and the lifecycle distribution.

Figure 14. DECOM Lifecycle Cost FY2015

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Appendix B Navigation and Communication System

The DECOM data handling system is shown in Figure 15. DECOM data handling system shows how all

of the DECOM subsystems’ data are sampled via their respective input busses and transferred to the

DMCs. These signals are then stored or sent to the Communication System for transfer to the Ground

Station via TDRSS. The DECOM main Computer will be a set of two VMEbus type computer systems

with SPARC V7 type single board processors that operate with a 32 bit system as a master/slave pair for

redundancy in order to handle the 3 Mbit/sec requirement.

Figure 15. DECOM data handling system

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Appendix C Power

Figure 16 shows how the UltraFlex solar arrays provide power to the Main Power Production Terminal

for distribution to the Bus Regulator and Charge Controller for battery charging and distribution via the

Voltage Regulator and Power Filter. The NEXT Engine, DRA, and Phased Array Amplifier each have a

dedicated bus to minimize electrical switching noise on the smaller 28VDC and 5VDC load busses which

power the spacecraft’s navigation, communication, temperature and pressure monitoring sensor systems.

Figure 16. DECOM Power Distribution Flow Diagram

Table 10. DECOM Power Budget by mission mode. Mission Mode Expected Power Loadings Power (W) Total (W)

Transfer (Daylight) Next Thruster 6900

Thruster Gimbaling 32

Thruster PPU 300

Battery Charging 4991

Phased Array Antenna 200

Radiators (2 m^2 estimated) 0 12423

Transfer (Eclipse) Next Thruster 6900

Thruster Gimbaling 32

Thruster PPU 300

Heaters (1 m^2 estimated) 500

Phased Array Antenna 200 7932

Observation Phased Array Antenna 200

Close Approach Camera 7

Navigation 22

Observation Camera 7

Radiators 0

Heaters 500

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Next Thruster 6900

Thruster Gimbaling 32

Thruster PPU 300

Floodlights 500

LIDAR 50 8518

Capture Robotic Arm 700

Floodlights 500

Tracker (Kinect-type device) 12

Navigation 22

Radiators 0

Heaters 500

Low Bit Camera 7

High Bit Camera 7

Arm Heaters/Radiators 50 1798

De-Spin ADCS 32

Patch Antenna 120

Sun Sensor 12 164

De-Orbit (Daylight) Next Thruster 6900

Thruster Gimbaling 32

Thruster PPU 300

Battery Charging 4991

PCDU 140

Phased Array Antenna 200

Radiators (2 m^2 estimated) 0

Robotic Arm 70

Floodlights 0

Tracker (Kinect-type device) 12

Navigation 22

Heaters 10

Force Sensors 20 12697

De-Orbit (Eclipse) Next Thruster 6900

Thruster Gimbaling 32

Thruster PPU 300

Phased Array Antenna 200

Radiators (2 m^2 estimated) 0

Robotic Arm 70

Floodlights 500

Tracker (Kinect-type device) 12

Navigation 22

Radiators 0

Heaters 500

Force Sensors 20 8556

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Appendix D Structures

Table 11. Mass Budget Subsystem/Component Unit Mass (kg) Quantity Total Mass (kg)

Payload

345.426

Hand 19 1 19

Control Arm 321.5 1 321.5

Camera - Close Range (CR) 0.258 1 0.258

CR Optics 0.1 1 0.1

Camera - Long Range (LR) 0.258 1 0.258

LR Optics 0.2 1 0.2

Video Recording/Compressor 1.11 1 1.11

Computer 2.5 1 1

Flood Lights 0 0 0

LIDAR - Optical Head Unit (OHU) 4 1 1

LIDAR - Avionics Unit (AVU) 6 0 1

Primary Propulsion

119.43

Main Engine 12.7 1 12.7

Low Pressure Array 3.1 1 3.1

High Pressure Array 1.9 1 1.9

Main Engine Gimbaling 6 1 6

Main Engine Tankage 20.41 3 61.23

Thruster PPU 34.5 1 34.5

ADCS

129.34

Monopropellant Thrusters 0.5 16 8

Momentum Wheels 10.5 4 42

Hydrazine Tankage 28.12 1 28.12

Pressurant Tankage 24.25 1 24.25

Regulator 1.72 1 1.72

Isolation Valve 0.36 25 9

IMU/Gyro 4.25 1 4.25

ADCS Fuel Lines 6 2 12

Power

238.9

Primary Solar Arrays 45 2 90

Maintenance Solar Arrays 5 1 5

Power Conditioning and Distribution Unit (PCDU) 44 1 44

Battery Packs 33.3 3 99.9

Navigation

20.632

Star Tracker 1.85 2 3.7

GPS 9.5 1 9.5

Sun Acquisition 0.054 8 0.432

Earth Acquisition 3.5 2 7

Comms & Data Handling

21.95

Phased Array 7 2 14

Patch Antenna 0.075 6 0.45

Computer 2.5 3 7.5

DECOM Mission

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Appendix E Thermal

The thermal subsystem must maintain the spacecraft internal and external components within their safe

operating temperatures at all times. Failure to do so may result in component or mission failure. A

thermal analysis of the spacecraft and components was done using the Systema satellite thermal analysis

software. The model gave the temperatures of all faces and components as functions of time for a

simplified spacecraft orbit.

Figure 17. Thermal Profile of DECOM

Surprisingly, even with internal power dissipation, internal radiative transfer and conduction, and

assuming a worst-case temperature environment, all internal components were within their operating

temperatures, although many were at high end of their safe values. However, actual temperatures of the

craft were expected to be higher because of several assumptions made in the model: resistive heating was

ignored, conduction from the thruster to the craft was neglected, spikes in Earth albedo and infrared

radiation were averaged out, and finally free molecular heating in the atmosphere was ignored. For these

reasons, a craft coating was chosen that reduced all component temperatures to lower portion of

temperature range. This passive coating was all that was needed to adequately maintain the temperatures

in the model, although in actual practice heaters would be used on propellant lines, and heat pipes would

likely be used on batteries to remove large heat generated during charge/discharge.

DECOM Mission

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Appendix F Transient Arm Analysis

The figure below is a screen capture from an animation which shows the simplified arm (pitch joints

only) simulation of capture of a rotating Cosmos satellite. The forces and torques in the joints are plotted

along with a visual representation of the arm geometry through capture. The curve colors in the plots

correspond to the joint colors in the arm.

Figure 18. DRA Arm Transient Analysis

DECOM Mission

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Appendix G Outreach Activity Lesson Plan and Picture

DECOM Mission

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DECOM Mission

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DECOM Mission

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University of Washington engineering students’ second visit

“Your CEV Design”

Student Handout Student name:

Lesson Objective The Space Shuttle is being retired this year and NASA wants a new vehicle to replace it, as well as do much

more. Using the Engineering Process you will design and construct a Crew Exploration Vehicle (CEV for short)

for NASA. Your vehicle should be reusable and capable of carrying astronauts to the Moon and Mars safely.

You will record your design process and results on this handout.

Part 1. Design Goal

You CEV design must:

be able to travel to the Moon and Mars – a roundtrip distance of 300 thousand to 130 million miles!

carry 6 astronauts safely – a roundtrip time of a couple days to 1.5 years!

carry and maintain several scientific experiments

allow astronauts to land on the Moon or Mars to explore and collect samples

Research Observations Answer the following questions by talking with your group and the University of Washington engineering

students:

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1) What do you think astronauts will need to survive in space such a long time?

2) What do you think the CEV needs to transport the astronauts and experiments?

Watch the “Inside the Space Shuttle” presentation and work with your group to answer the following question:

3) What are the key parts of the Space Shuttle?

Part 2. Design Plan

4) Develop a design for your CEV by filling out the design table below:

5) Next, use a piece of graph paper and to draw your CEV, labeling where your components are located, and

including space for the astronauts to move around in. Discuss your design with the University of

Washington students.

Construction Now you will work with your group and the University of Washington engineering students to construct a CEV

model following your group’s designs.

6) List in the table below the problems you ran into during this process and how you solved them:

7) Draw your final constructed CEV design in the space below and label the components:

Part 3. Conclusions:

8) What was the hardest part about this activity?

9) How long do you think it will take NASA engineers to use this engineering process to build a CEV?

10) What was your favorite part about this activity?

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