Design and Development of a CubeSat De-Orbit Device

faithfulparsleySoftware and s/w Development

Nov 2, 2013 (3 years and 8 months ago)

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1

Design and Development of a CubeSat De
-
Orbit Device








Students:

An Kim

Cian Banco

David Warner

Edwin Billips

Langston Lewis

Thomas Work

Mackenzie Webb

Benjamin Crawse

Jason
Harris





Faculty Advisor:

Dr. Robert Ash








2

Abstract



CubeSats, shown in Figure A
, are a common satellite used by universities and
institutions for research in space. At 1000 cm
3

the CubeSat can be integrated into many
different rocket payloads. A
CubeSat satellite will orbit at 900 km altitude from the earth
for 1,000 years without any type of brake or modifications for it to de
-
orbit in less time. It
is the objective of our team to design and test a prototype of an aerodynamic brake that
can be pl
aced on a CubeSat and effectively reduce its time in low earth orbit to within
1000 years. It is also our objective to design and test a prototype of a deployment
device that can be mounted and eject our CubeSat form a sounding rocket payload. It is
our go
al to define the telemetry necessary for the mission of our CubeSat from the
launching of the sounding rocket, to the timing and ejection of our CubeSat and
deployment of the aerodynamic brake. Telemetry will also be needed for the verification
of deployme
nt of the brake and adequate drag.






Figure A 1U CubeSat









3

Table of Contents

Abstract..............................................................................................
...............................2

Table of
Contents.....................
....................................................................................
....
3

1.

List of Figures....................................................................................
...........................
4

2.
Introduction.........................................................................................................
..........
5

3.
Sounding Rocket Test Flight..............................................................
...........................
7

4.
Deployment Device.....................................................................................................
10


4.1 Introduction………………………………………………………………………..10


4.2 Previous Research……………………………………………………………….11


4.3 Conclusion………………………………………………………
………………...12

5.
De
-
orbit Device...........................................................................................................
13


5.1 Introduction………………………………………………………………………..13


5.2 Previous Research………………………………………………………………..13


5.3

Conclusion…
………………………………………………………………………17

6.
Electrical Circuitry: Micro
-
Controllers and Various, Necessary Components

............18


6.1 Introduction………………………………………………………………………..18

6.2 Analysis……………………………………………………………………………18

6.3 Conclusion/Future Work………………………………………
………………….19

7.
References........................................................................................
.........................2
1

8.
Updated Gantt Chart..........................................................................
....................
.....
22

9. Conclusion
…………………………………………………………………………………..23
























4


1.

List of Figures

i)

Figure A 1U CubeSat

ii)

Figure 3.1 RockSat
-
X deck plate

iii)

Figure 3.2 Sounding Rocket Payload Stack

iv)

Fig 5.1a Orbiting height as a function of time without a
de
-
orbiting device

v)

Fig1.1b Orbiting height as a function of time with a de
-
orbiting device








































5



2. Introduction


Since the beginning of the Space Age, thousands of defunct satellites,
components, and associated debris have
accumulated in Earth’s immediate orbital
neighborhood. Orbiting at high altitudes, their lifetimes can span a thousand years or
more, and during that time will remain a hazard to all satellites with similar orbits. The
hazard of space debris was made appa
rent when the windshield of the Space Shuttle
was damaged significantly in a collision with a tiny flake of paint that had a relative
velocity of 2 km/s to 8 km/s.
[1]

Collectively, this group of orbital refuse is known as space
debris, or “space junk”.

Though no treaty exists for the mitigation of space debris, the United Nations has
formed an eleven
-
nation study to address the issue, and the United Nations Office for
Outer Space Affairs has set up a group of mandates meant to stop the growing space
debr
is problem. The Inter
-
Agency Debris Coordination Committee (IDAC) has pressed
NASA, ESA, and private groups to deorbit satellites orbiting below 2000km within 25
years of mission completion. Most picosatellites (CubeSats) fall into this area, as a
typical
orbit is only at a 900 km altitude.

CubeSats are often launched as hosted
payload aboard a bigger rocket, and serve as “bricks” to meet payload weight
requirements. At a 900 km altitude, a CubeSat can remain in orbit for well up to 1,000
years, and may pr
ove to be a hazard to other satellites, present and future.

Our goal is to develop an aerodynamic brake which will diminish a satellite’s
orbital time from 1,000 years to the mandated 25 years. An aerodynamic brake will take




6

advantage of the conditions of
near
-
Earth space, which is not a perfect vacuum; enough
monatomic oxygen is present to provide a noticeable effect of drag on a structure with a
large surface area. It is this same effect that deorbited Skylab earlier than expected
(before it re
-
entered, a

rendezvous with the then
-
new Shuttle was planned), and which
now requires that the International Space Station be periodically boosted into higher
orbit. Echo 1, a passive communications satellite launched into Earth orbit, experienced
a measurable drag e
ffect, largely in part to its 40m diameter balloon, which provided a
large surface area for drag in orbit.

We propose to develop an aero
dynamic
brake component small enough to be
installed on commercial CubeSats, using only a minimum of available volume an
d mass
constrained by the CubeSat design, and with off
-
the
-
shelf constituent parts. Using a
single
-
unit (1U) CubeSat (10cm x 10cm x 10cm) as our mission template, we will
pursue two main options to reducing a satellite’s lifetime. One option is to use
sub
limating benzoic acid to inflate a 1.22m diameter balloon; benzoic acid goes from
solid to gas at low pressures and mild temperatures, making it attractive for application
in a space environment. The inflated balloon would present a large surface area for
drag to take effect. The alternative option is to use Shape Memory Alloy (SMA) to
deploy a similar balloon
-
shaped structure. Nitinol, a commercially available
SMA,

expands and contracts under heat, making it able to flex or draw out complex
structures. By
evenly heating a Nitinol wire assembly, the aerodynamic brake can
deploy in full. The option of using benzoic acid along with Nitinol wire to reinforce the
shape of the aerodynamic brake will also be explored.





7

We plan to test both deployment options until
one is clearly superior. In addition,
we will examine our material options for the balloon structure; we are looking at Kapton,
Mylar (PET film), and aluminized versions of said materials for the balloon structure. We
plan to test the balloons in a vacuum
chamber, and validate the aerodynamic brake’s
functionality with a suborbital flight. The ultimate objective of the project is to design and
test a de
-
orbit system for CubeSats that can reduce its time in low earth orbit to within
25 years. The de
-
orbit de
vice must prove to be a robust and viable system having
commercial viability at a minimum cost. Three current options for performing the
experimental tests necessary to accomplish these objectives are through a high altitude
balloon, orbital demonstrator,
and suborbital sounding rocket. A suborbital sounding
rocket flight is the current option being pursued by the team.



3. Sounding Rocket Test Flight



The current platform for the necessary experimental test of our CubeSat with
an

installed de
-
orbit devic
e currently being pursued by the team is through a sounding
rocket test flight. Sounding rockets provide a relatively simple and inexpensive way for
students and researchers to perform research in a low earth orbit environment.
Sounding rockets are divided

into two parts: the payload and a rocket motor.
[3]

When
the sounding rocket is launched it takes a parabolic trajectory. As the rocket motor uses
its fuel it separates and falls back to earth while the experiments placed on the payload
are conducted in th
e low earth orbit environment. Once the payload re
-
enters its gently
brought down to earth by a parachute.





8


The program the team is currently pursuing for a sounding rocket test flight is
NASA’s RockSat
-
X program at Wallops Flight Facility. The RockSat
-
X p
rogram utilizes
the Terrier
-
Improved Malamute Suborbital sounding rocket that will reach apogee at
approximately 160 km altitude from earth.
[2]
The sounding rocket’s skin and nose cone
will eject after second stage burn
-
out of the Malamute launcher. RockSa
t
-
X
utilizes
decks shown in Figure 3
.1 for customers to mount their payloads as opposed to the
canisters used in its predecessor RockSat
-
C. The RockSat
-
X decks are in a stacked
configuration interconnected by longerons spanning the entire length of the Roc
kSa
t
-
X
payload as shown in Figure 3
.2. The decks are designed to provide customers direct
access to the environment of space and allow payloads to deploy booms and other
mechanical devices once the skin is ejected.






9


Figure 3.1 RockSat
-
X deck plate


Payload Stack













Figure 3.2 Sounding Rocket Payload Stack






10

Wallops Flight Facility still provides sounding rocket flights through the
predecessor of Roc
kSat
-
X; the aforementioned RockSat
-
C program. The team
considered both options. due particularly to the difference in cost to participate in the
program. The RockSat
-
C is the cheapest option at $12,000 for a canister compared to
$24,000 for a RockSat
-
X dec
k plate. The access that RockSat
-
X provides to the low
earth orbital space environment that will allow for the ejection of a CubeSat, along with
the superior power and telemetry provided compared to RockSat
-
C are the primary
reasons that it is determined b
y the team to be the best sounding rocket option for our
experiment. The team intends to submit an “Intent to Fly” form with the necessary
$2000 deposit to Wallops for the RockSat
-
X sounding rocket launch in 2014.




4. Deployment Device


4.1 Introduction

The team has designed the “ODU Picosatellite Orbital Deployer”. This design is based
on California Polytechnic State University’s “Poly Picosatellite Orbital Deployer”, or P
-
POD for short. This design has become the standard for many successful orbital fli
ghts.
Our team’s design, the “O
-
POD”, plays a critical part in getting the cubesat away from
the sounding rocket. The O
-
POD houses a spring inside its length, which is compressed
while the cubesat is stored within. A quick release mechanism holds the sprin
g and
cubesat back, until a timer event from the RockSat
-
X power interface signals its release.




11

With the deployer freed, the spring pushes the cubesat out at a modest lateral velocity,
1.6 m/s. This imparted motion gives the cubesat clearance to operate th
e De
-
Orbit
Device, away from the rocket.

Although the cubesat is constrained severely by mass, size, and power requirements,
the O
-
POD rests on a deck which has room for 30 lb mass, and 1 amp
-
hour of provided
electrical charge. This deck rotates with the s
ounding rocket at 0.5 Hz for the majority of
the microgravity window, and is directly exposed to the vacuum conditions of space.
Unlike the cubesat, the O
-
POD remains in the sounding rocket, and can be considered
a recoverable, provided thermal insulation
for wires. The team has considered sharing
pins for the deck’s power interface with possible flight partners, in the interest of
reducing cost. As only one timer event is required for the O
-
POD, this opens up options.

4.2 Previous Research

The previous tea
m assessed the following design criteria when making the O
-
POD:


1.

Must withstand 25g loads

2.

Minimal weight & material

3.

Eject cubesat at ~1.6 m/s

4.

Must withstand vibration tests

The dimensions of the O
-
POD, in Figure 4.1 are 12.7cm W x 12.7cm H x 19.62cm L,
the same dimensions as the span on the RockSat
-
X deck plate. The structure is made
of bolted Al
-
7075
-
T651 plates, a mate
rial which is high strength and can easily be




12

machined
.
When joined, the plates form a rectangular box with internal rails, which
provide a guideline for a pusher
-
plate and the spring. The plates are 7mm thick, and the
entire assembly is 1.477kg.

Extensive PATRAN analysis has been performed on the O
-
POD, throug
h five iterations
of design. Each step refined the model’s mesh and geometry, providing increasingly
accurate results with the improvements.

The previous team worked on “pinned” measurements of ejection velocity, using
distance markers, manual release, and

a ca
rdboard mockup of the CubeS
at. This
mockup’s coefficient of friction varied with humidity, as it was constructed of cardboard.
Tests were performed without deciding on a quick release mechanism option, as
choosing was deemed premature. However, linear

solenoids were considered.


4.3
Conclusion/Future Work

The present team is working on continuing the PATRAN analysis started by the
previous team, with the intent being to account for a mounting plate between the O
-
POD and the RockSat
-
X deck. Velocity mea
surements with a newer mockup, of
Aluminum construction, will provide more accurate results than a cardboard one. The
current team plans to choose a quick release mechanism option as soon as possible,
with consideration going toward Vectran Line Cutters or

Linear Solenoids, as suggested
by the previous team for its budget
-
conscious and simple design advantages.









13

5. De
-
Orbit Device


5.1 Introduction



The objective of the de
-
orbit device is to successfully deploy a spherical balloon
from a 1U
CubeSat 9x9x1 cm storage space. When the spherical balloon is fully
inflated, the corresponding drag will act like an aerodynamic brake that will greatly
reduce the 1U CubeSat de
-
orbiting time. The main concerns of the de
-
orbiting device
after deployment
are the ability to withstand ultraviolet radiation, atomic oxygen,
changes in temperature, and micro meteor impact.



Weight and volume are critical constraints of the de
-
orbiting system due to the
restrictions of the CubeSat. Sending payloads int
o orbit is based on cost per unit mass,
and varies by space programs. Therefore, CubeSat must be small in volume and mass
but must have sufficient space to fit the de
-
orbit mechanism. Since the CubeSat
provides low budget access to space, the de
-
orbit dev
ice must also be economical.
Overall, if the CubeSat de
-
orbit mechanism is intended to be used commercially; it must
be small, affordable, and able to withstand the conditions of space over vast periods of
time.


5.2

Previous Research

The starting point

of the CubeSat project was to research and design a suitable
membrane for the de
-
orbiting mechanism. The study focused on the membrane portion
of the de
-
orbiting mechanism through the lifetime of a 1U CubeSat. A prototype may be
modeled for the de
-
orbit
mechanism based on the study.





14



A proper polymer for constructing the membrane must withstand ultraviolet
radiation and atomic oxygen; and the best available material is Upilex® brand polyimide
film produced by UBE industries. Upilex® is commonly us
ed to protect against
ultraviolet radiation and can also protect against atomic oxygen with the addition of a
silicon oxide coating.



Another key aspect of the aerodynamic brake mechanism is the shape and size
of the membrane. When selecting shape
s of the membrane, the shape must meet a
certain cross
-
sectional area to produce a desired drag. Based upon assumption, the
membrane cross section will always be facing perpendicular to the direction of travel.
However, this assumption is not entirely accu
rate. The CubeSat will rotate, turn, and
tumble. The position of the membrane may change based on the orientation of the
CubeSat and may not yield its maximum cross
-
section projected into the direction of
travel. To combat this potential problem, we will c
onstruct the membrane in the shape of
a sphere.



Previous groups also conducted research using ODU’s virtual lab Satellite Tool
Kit (STK) simulation and concluded it would take the 1U CubeSat at least 480 years to
return back into Earth from an alt
itude of 800 km under unassisted conditions (Figure
1.1a). The National Space Treaty requires all satellites that accomplish their missions
to descend back to Earth in under 25 years. Through the STK simulation, it was
determined that the 1U CubeSat satel
lite would return to Earth within 126 months with a
minimum cross
-
section area of 0.5625 square meters (Figure 1.1b). This would allow
the U1 CubeSat to descend without violating the National Space Treaty.






15


Fig 5.1a

Orbiting height as a function of time

without a de
-
orbiting device



Fig1.1b

Orbiting height as a function of time with a de
-
orbiting device







16

Another great source of background info
rmation for the aerodynamic brake is the
NASA’s Echo I Satellite project. In 1960, NASA deployed a 100
-
foot diameter inflatable
satellite using a mix of sublimating compounds. NASA used benzoic acid and
anthraquinone to reduce the need of compressed gas
, thus reducing the need of
regulators and storage device. The benzoic acid was evenly distributed within the
membrane of the satellite. When mixed with anthraquinone, the benzoic acid turned
from a fine solid powder into a paste, allowing the compound to

inflate the satellite. The
sublimation process of benzoic acid required minimum solar heat to inflate, but
problems arise as the satellite passes through the shadow of Earth (Clemmons).

Through the successful launch of NASA’s Echo 1 Satellite, ODU’s form
er team
tried to apply the same deployment method towards the CubeSat de
-
orbit mechanism.
Though NASA used benzoic acid and anthraquinone, the team was restricted from
using anthraquinone due to financial limit. Even without anthraquinone, benzoic acid
al
one at 900 km altitude with solar heat will produce a vapor pressure of between 3 to 5
torr (400~650 Pa). This estimation of pressure will be used to inflate the de
-
orbit
mechanism as the exterior pressure at 900km altitude is minimal.

Another process the

ODU former team carried out was tests of benzoic acid’s
inflation with the use of store bought Mylar balloons. The sizes of the Mylar balloons
were 18 inches in diameter; one balloon was filled with 10 grams and the other with 15
grams or acid. Air was
removed from the balloons and sealed with space grade epoxy
to ensure sealant in a space
-
like environment. The balloons were then folded tightly
and tested within a vacuum chamber under ambient room temperature. The test




17

inflations were successful; howev
er no record of Mylar balloon fully inflated within the
300 second limits.


5.3 Conclusion/Future Work

Based upon the successful inflation tests made by the previous ODU CubeSat
team, benzoic acid is currently the best choice of inflating the U1 CubeSat de
-
orbiting
system. The next objective of the inflation tests is to determine the temperature at
which the benzoic acid will inflate the aero
dynamic
brake within the restricted time.
According to Emerald Performance Material sheet, benzoic acid will sublim
ate at
roughly 12.5*C at 160 km altitude. In order to achieve full inflation of benzoic acid in
less than 300 seconds, a thermistor will be added to the test system to increase the
system’s temperature.

The second goal of the inflation tests is to verify
the amount of benzoic acid to
inflate the 18 inches balloons. The mass of benzoic acid must be calculated and evenly
distributed within the Mylar balloon. The even distribution of benzoic acid will assist in
the unraveling of the tightly folded balloon.

A
lthough our test flight will only reach an apogee of 160 km, we will be able to
use data acquired from the flight to use toward commercial production. Upon a
successful mission with the RockSat program, our team will have a verified model upon
which to bas
e our future calculations for commercial use upwards of 900 km. While
there may be some complications that may be unforeseen, we remain confident that we
will be able to overcome any difficulties.






18


6
. Electrical Circuitry: Micro
-
Controllers and Various, N
ecessary Components


6
.1 Introduction

The successful deployment and verification of the CubeSat aero brake will
depend upon several electrical components working seamlessly together. In order to
accomplish this task it will be necessary to provide power a
nd communication onboard,
allowing commands to be sent from wallops to the CubeSat. These commands will then
drive the release mechanism and heating element for the aero brake. Once inflated, it
will also be necessary to provide verification of the aero
brake inflation. The CubeSat
electrical team has recruited the help of Benjamin Cawrse (CS major) and Jason Harris
(TCC) to reach these goals.


6
.2 Analysis

These objectives will require the integration of various components into a micro
-
controller inside

the CubeSat. The electrical team has already come into possession of
the desired micro
-
controller, an Arduino Uno. The board has 14 pins for digital input
and 6 for analog input which is adequate for our goals. As open source software, the
programming i
nterface can be downloaded directly from the official Arduino website.
Coupled with an active community and vast code libraries, a large variety of devices can
be connected to the board. As a micro
-
controller capable of supporting a wide
assortment of de
vices, from accelerometers to video acquisition, it suits our purposes
fully.





1
9

The CubeSat electrical team is considering the usage of space rated lithium
batteries to provide power to the board. Shipped with the Arduino were two 9
-
volt
batteries and the n
ecessary wiring to attach it to the board. While these were sufficient
to conduct simple testing, it will be necessary to have a larger current capacity in order
to run the various components long enough for successful deployment.

For video acquisition an
d radio communication, the CubeSat electrical team is
considering the usage of camera that will repeatedly take pictures throughout the
inflation cycle at a resolution of 160x120. The Arduino on the rocket deck will connect
up to the telemetry interface on

the rocket to relay information at 19,200 baud using the
XBee wi
-
fi transmitter
[2]

.The Arduino on the CubeSat will relay information from a
camera and some sensors back to the rocket deck. The video data will be stored on a
SD card on the rocket deck for

later retrieval and the sensor data will be relayed real
time back to the operations center via the telemetry interface. The options for video and
communications are both low
-
power and low
-
weight, making them ideal for this project.

As of October 11th, 2
012, the CubeSat electrical team has access to the
vibrations lab, overseen by Dr. Bilgen. This lab provides with strain gauges which will
be used to monitor the drag produced by a fully inflatable balloon. Provided with a small
current and using a resis
tor configuration known as a wheatstone bridge, the changes in
voltage can be converted into a related strain value, from which full inflation of the
balloon can be determined. Built into the Arduino board are accelerometers, which can
also monitor the gr
adual deceleration of the CubeSat.

In order to provide sublimation within an acceptable timeframe, a NTC (negative
temperature coefficient) thermistor will be used to provide the necessary heat to




20

catalyze the reaction. As a special kind of resistor, it w
ill be able to generate the
required heat when minimal current is applied.



6
.3 Conclusion/Future Work

Currently, the team is working towards the acquisition and integration of the
specified components into the Arduino board. With every compon
ent being fully
compatible with the Arduino, there should be no issues with connectability. We are on
schedule for this task and aim to have it completed in order to advance to vacuum
chamber testing. Once communication can established with the CubeSat a
ero brake,
instructions can be sent remotely during vacuum chamber testing, giving great insight
into the behavior of the system, especially with regards to data retrieval.


























21

7.
References


[1] K. S. Edelstein, "Orbital Impacts and
the Space Shuttle Windshield," E. S. a. D.
Branch, Ed., ed. Houston, TX: NASA, 1994.


[2] Colorado Space Grant Consortium (2013).
ROCKSAT
-
X Users’ Guide
(Revision 3)
[Online]. Available:
http://spacegrant.colorado.edu/COSGC_Projects/RockSat
-
X%202013/RockSat
-
XUsersGuide2013Rev3.pdf


[3] E. M. Marconi. (2004). What is a Sounding Ro
cket? Available:
http://www.nasa.gov/missions/research/f_sounding.html


[4] Benzoic Acid, PRODUCT INFORMATION BULLETIN
http://www.emeraldmaterials.com/epm/kalama/micms_doc_admin.display?p_customer=
FISKALAMA&p_name=PRODBULL
-
BENZOICACID.PDF, Emerald Kalama C
hemical,
LLC, 2012.


[5] Scheider, Kimberly. “
Parametric Study of Solar Activity and Aerodynamic Drag
Coefficient Versus Orbital Decay Period
.” STK Simulation. Old Dominion University,
2012. Print.


[6] Clemmons, Dewey L. Jr. "The Echo 1 Inflation System."

NASA Technical Note.
NASA Langley Research Center, 1964. Print.






22

8.
Current Gantt Chart






23

9. Conclusion



During the first half of th
is semester
the team has spent the majority
of its time
gathering material
s, acquiring

information and planning. The information and resources
will be utilized for performing experimental tests and creating prototypes

necessary

for
the CubeSat, aerodynamic brake and deployment device. These tasks will be executed
to interface with RockSat
-
X
. Although other options will be considered by the team the
successful completion of our objective of testing and verifying a de
-
orbit device on a
CubeSat in space.