Scalable PnP Drag Sail Module Deorbit System for LEO Satellites

Abstract

The increase in use of Low Earth Orbits (orbits typically considered ranging from 250km to 2000km) for research, student, and global constellation satellites has increased the satellite density by large, posing a threat of collisions with other satellites and debris satellites. To avoid this, satellites are recommended to be deorbited as soon as possible after their mission lifetime. Various deorbiting techniques are being worked upon, especially passive ones as their size, mass and power budgets are usually lower than that of active ones. Drag sails are one such passive method popular in the field. In this paper, we propose a plug and play design of a drag sail module, different from the mechanical boom and electrical methods currently used, using Commercial-Off-The-Shelf (COTS) components for cubesats and satellites up to 500kg in mass for LEO orbits with altitude up to 1100km. A conceptual design is proposed, keeping the mission requirements from standard sample missions and trajectory analysis results in focus. This is followed by a conceptual control system analysis and cost analysis. The research provides a ready to use, scalable solution for the mentioned problem with COTS components and standard processes for trustworthy acquisition.

Full text

Scalable PnP Drag Sail Module Deorbit System for LEO
Satellites
1Anshuman Shukla, 1Pranav Milind Sawant and 2K.C. Mohite
1Student, C.P.V.N Kaimganj, U.P.,
India;
1Student, Army Public School, Pune,
India;
2Principal, CT Bora College,
Savitribai Phule Pune University,
Pune, India;
Email:
{1anshuman.contact07@gmail.com,
1pranavsawant132@gmail.com,
2kcmohite@gmail.com}
Abstract
The increase in use of Low Earth Orbits (orbits typically considered ranging from 250km to 2000km) for research, student, and global
constellation satellites has increased the satellite density by large, posing a threat of collisions with other satellites and debris satellites. To avoid
this, satellites are recommended to be deorbited as soon as possible after their mission lifetime. Various deorbiting techniques are being worked
upon, especially passive ones as their size, mass and power budgets are usually lower than that of active ones. Drag sails are one such passive
method popular in the field. In this paper, we propose a plug and play design of a drag sail module, different from the mechanical boom and
electrical methods currently used, using Commercial-Off-The-Shelf (COTS) components for cubesats and satellites up to 500kg in mass for LEO
orbits with altitude up to 1100km. A conceptual design is proposed, keeping the mission requirements from standard sample missions and
trajectory analysis results in focus. This is followed by a conceptual control system analysis and cost analysis. The research provides a ready to
use, scalable solution for the mentioned problem with COTS components and standard processes for trustworthy acquisition.
Keywords
Drag Sail, Satellite Deorbiting, Passive Deorbiting Systems, Sail Deployment, Plug and Play Design
Abbreviations
LEO - Low Earth Orbits
PnP - Plug and Play
COTS - Commercial-Off-The-Shelf
ADM - AirDragMod
IADC - Inter-Agency Space Debris Coordination Committee
UTIAS SFL - University of Toronto Institute for Aerospace Studies Space Flight Laboratory
ATOX - Atomic Oxygen
UV – Ultraviolet
WSCEA - Whole Solar Cycle Effective Area
ESA- European Space Agency

1. Introduction
In recent years, Low Earth Orbit (LEO) satellites have
been constantly facing the problem of space debris
requiring them to perform collision avoidance
maneuvers [1]. With the growing consciousness for
space debris, it is only a matter of time before satellite
makers are compelled to incorporate a deorbiting
system.
Heavier satellites could carry some extra fuel to
perform deorbiting burns but are constrained by mass.
Cubesats, due to their small size, don’t face this often
but with their growing popularity, it is important to
ensure their timely deorbit. A collision avoidance
maneuver is almost impossible for cubesats lacking any
active control system owing to its limited mass, size
and cost constraints, which reinforces the need for such
a passive deorbiting system in order to avoid such
scenarios in future.
Since deorbiting isn’t an integral phase of a mission
and doesn’t affect its operations directly, a deorbiting
system would have to be as “invisible” as possible viz.
low mass, low volume, low operational power usage.
Although the launch costs have significantly fallen [2],
mass still affects satellites power usage during
necessary attitude control thus affecting mission life.
The two most discussed and worked upon passive
techniques for deorbiting are drag sail and plasma
brakes. While plasma brakes are good for cubesats in
LEO orbits, they aren’t found to be equally effective in
upper orbits for satellites of higher mass [3]. So, in this
research paper, we propose a drag sail design which is
easily scalable, easy-to-implement and uses low power.
A drag sail deorbiting system is a system which uses a
thin sail to maximize drag area of a satellite thus
increasing perturbations of aerodynamic drag force.
𝐹𝑑 = 1
2𝜌𝐴𝐶𝑑𝑣2 (1)
Equation (1) is a basic equation governing aerodynamic
drag on a surface. Here, 𝐹𝑑 is the drag force, 𝜌 is the
density of the medium, 𝐴 is the cross-sectional area i.e.,
area perpendicular to atmospheric flow vector, 𝐶𝑑 is the
coefficient of drag generally taken 2.2 for LEO orbits,
and 𝑣 is the orbital velocity of the craft. While the
atmosphere of Earth grows thinner as higher orbits are
attained, it still causes significant degradations in orbit
of satellites in LEO over time which need to be
accounted for using boost burns. A drag sail increases
the rate of orbital decay, in compliance with current
Space Debris Mitigation best practices and guidelines
[4].
Although not a new concept, significant work has
only been done in the past decade with TechEdSat4exo-
brake launched in 2014 and successfully deorbited with
a 0.35m2 of drag sail area. TechDemoSat-1 launched in
2014 with 6.2m2 of drag sail area, dragNETTM - a
mission which successfully deorbited a Minotaur upper
stage in 2016 utilizing a 14m2 drag sail area, CANX-7
(Canadian Advanced Nanosatellite eXperiment-7) [5]
in 2017 with a 5m2 of drag sail area, removeDebris in
2018 with a drag sail area of 0.35m2. There are more
missions currently in orbit or to be launched including
the ADEO-N2 subsystem of 1U (10cm x 10cm x 10cm)
size capable of stowage of 5m2 of drag sail by ESA’s
Clean Green Space Initiative. [6]
In this paper, we propose a Plug and Play (PnP)
module design standardized using Commercial-Off-
The-Shelf (COTS) components as a ready to use
solution for the purpose of deorbiting. PnP design
philosophy enables quick and reliable installation. This
deorbiting module has been designed in two
configurations: one for cubesats and other for heavier
satellites of up to ~400kg in an orbit of up to ~900km.
The cubesat configuration has two further variations.
The design is inspired by JAXA's IKAROS [7] mission
which was the first interplanetary solar sail propulsion
mission launched in 2012 and performed a flyby of
Venus. It weighed 300kg and consisted of a 14m x 14m
spinning sail. It used angular momentum leftover after
separation from its rocket stage to deploy 4 tiny masses,
which due to centrifugal force extended deploying the
sail attached to it. This technique is much more efficient
than electrically or stored mechanical energy deployed
systems as we would further discuss in this paper.
Being aware of all constraints, this work proposes a
drag sail module- AirDragMod (ADM) design to
deorbit cubesats and satellites of mass up to ~400km up
to an orbital altitude of ~900km. Thus, in order to
justify the capability of the design, a trajectory analysis
of simulations ran in FreeFlyer® has been performed,
keeping in mind the results, the design has been
proposed and is divided into three phases: i. Conceptual
design of both configurations, ii. Sizing of drag sail and
module dimensions are calculated, iii. Control System
conceptual analysis Then a comparison and cost
analysis has been shown, and lastly concluded with the
scope of further research on this.
2. Requirement Analysis
In order to determine the design requirements of
AirDragMod, two different reference missions have
been chosen, one for each configuration. For the

cubesat configuration, CANX-7 [5] is the chosen
mission, as it was a successful mission, the data from
which is publicly available. CANX-7’s main objective
was to demonstrate the use of an external drag sail
module. It was launched in a Sun Synchronous Orbit
(SSO) with an altitude of 688km and an inclination of
98°. It had an ADS-B receiver as its payload. The 3U
(10 cm x 10 cm x 34 cm) Cubesat platform weighed
3.75 kg and relied on a purely magnetic attitude
determination and control system. 2U was allocated to
space- craft housekeeping, payload systems and 1U
was allocated to its 4 dragsail modules stacked in pairs
of two one above another. The objective of the dragsail
module is to assist with the quickest deorbit possible in
coherence with the upper limit as set by Inter-Agency
Space Debris Coordination Committee (IADC)
guidelines which is 25 years [4]. In actuality, the
deorbit time would be much less. The dragsail was
deployed in 2017. It is expected to deorbit in roughly 5
years. This would later be used to derive the design
dimensions.
Considering these characteristics of the CANX-7
mission, the primary objective of ADM cubesat
configuration would be to deorbit a cubesat up to 10U
from LEO (up to 1100 km) orbit to below 100 km. In
addition, it would contribute to development of COTS
Plug and Play (PnP) passive deorbiting system for
cubesats and the advancement of dragsail induced
deorbiting technology. For this, a size constraint of 1U
(10 cm x 10 cm x 10 cm) form factor is set which
implies a mass constraint of 1 kg. This configuration
design must also be compliant with the CubeSat design
specifications and secondary payload launching
systems. For easy development and to reduce
development cost and time, space-tested components
and materials must be used and is a requirement.
Because of creating a PnP module, other functional
requirements arise which include: i. The ADM module
should be easily attachable and its dimensions
adaptable to host satellites. ii. Shall be independently
deployable and act as a completely independent system
using minimum host satellite power budget and
components.
It should, as mentioned, be capable of deorbiting
well within IADC guidelines. It should also be pointed
out that while CANX-7 has been used as a reference
mission for the cubesat configuration of ADM, it is not
limited to the CANX-7 type of mission. This
configuration in itself is intended to demonstrate
adaptability to other UTIAS SFL spacecraft such as the
Generic Nanosatellite Bus (GNB) and Nanosatellite for
Earth Monitoring and Observation (NEMO) bus
designs. An important requirement is the testability of
the design in a 1g environment or another controlled
environment easily creatable. [8]
The second configuration applies for satellites
greater than 10U (1m x 1m x 1m) in volume and greater
than 10kg in mass. Since there are no flown missions
of satellites greater than the above-mentioned
specifications which successfully deorbited using a
drag sail in authors’ knowledge, a result from ‘ADEO
De-Risk Dynamic Analysis’ or ADDA [9] which was
used by ESA for ADEO-N mission for the purpose of
de-risk activity analysis ADDA following directly the
ADEO (Architectural Design and Testing of a De-
orbiting Subsystem) activity, is used for reference. Its
purpose was the verification of the feasibility of de-
orbiting a spacecraft using a drag sail. One of its finding
was that the drag sail shortens the post mission lifetime
significantly, for example, a 25𝑚2 sail on a 300kg
satellite from a 600km orbit de-orbits the satellite 97%
faster in 5 years instead of 140 years. This would later
be used to compare simulations.
To sum up, the requirements for the second
configuration of ADM are mostly similar to that of the
cubesat configuration. The objective of this
configuration would be to deorbit a satellite of mass up
to 500kg in up-to an orbit of altitude 1000km. For this,
a size constraint of 2U (20 cm x 10 cm x 10 cm) is set
which also implies a mass constraint of 2kg. It is
ensured to be made of COTS components which are
space tested for easy development and to reduce
development cost and time. Similar to the cubesat
configuration, it has other functional requirements due
to being a PnP configuration viz. i. The ADM module
should be easily attachable and its dimensions
adaptable to host satellites. ii. Shall be independently
deployable and act as a completely independent system
using minimum host satellite power budget and
components.
Both the configurations’ system and their materials
must also withstand the harsh environmental conditions
in space such as the ATOX, UV and temperature
environment.
3. Optional Increase In Deorbit Rate
To increase the rate of deorbit by ADM simulations
were performed. The results are combinations of fixed-
attitude and variable-attitude simulations. The varying
attitude simulation results are needed as the fixed-
attitude simulations fail to account for the time varying
nature of the spacecraft's projected drag area. This
could be accounted for using an active control system
which is not viable for the lifetime of de-orbit.

Aerodynamic perturbations are not accounted for
below 450km, nor is geomagnetic disturbance due to
residual magnetic dipole moments above 650 km.
Assuming a constant drag area could yield very
misleading results and thus fixed-attitude simulation
results are needed.
To assess this, short term attitude simulations were
used to evaluate the de-orbit performance in slices of
an entire ~11 year solar cycle and combined
appropriately to arrive at a result which approximates
the performances. The output (WSCEA [5]) can be
compared against the constant area simulations
performed on FreeFlyerⓇ. The results in fig. 1 were
obtained on two configurations of 0-degree angle
between drag sailplane and residual dipole for 5 year
deorbit profile and 90 degree for 10 year deorbit profile.
The orbit considered was 700km altitude, 98.1-degree
inclination orbit with LTAN ranging 500 HHMM to
2800 HHMM.
These results indicate that providing occasional
stabilization in the form of nadir pointing/PID control
to keep drag area constant would increase the rate of
deorbit rate and thus decrease deorbiting time. The
most effective stabilization being at ~1000 HHMM
LTAN and ~2000 HHMM LTAN, as per the results,
thus maximizing the duration of higher drag area.
Similarly, stabilizations could be performed following
simulations of respective missions using the ADM in
future.
Figure 1: WSCEA results for mentioned orbit and inclinations
4. Adm Design
This section presents the design of the AirDragMod
(ADM) which is the core of the paper which was
carefully thought through several iterations. The main
stages can be distinguished: the conceptual design, the
deployment mechanism and design overview. These
phases are briefly introduced in the following
subsections.
4.1 Conceptual Design
During the conceptualisation of ADM’s design, the
selection of main configuration elements is performed.
For that figure of merits defined from mission analysis
defined from the mission analysis section is used. For
ADM, three main elements studied are: the drag sail,
the rotator, and the ADM structure layout. The design
criteria followed during the decision-making process
has been influenced by six key aspects: 1. fulfillment of
deorbit requirements, 2. better approach to deployment
than existing methods, 3. maximum drag sail area while
minimum ADM volume occupied, 4. generation of a
PnP module, 5. Preference for COTS components and
space tested technology, 6. manufacturable design at
minimum cost.
The ADM has been designed in accordance with the
results. One of the objectives of this paper is to propose
a scalable PnP module solving problems existing with

the current techniques of deployment, specifically the
stored mechanical energy boom deployment technique
used in CANX-7. The ADM is based upon the Solar
sail model used in the IKAROS mission by JAXA. The
idea behind the deployment was that angular
momentum of the module would cause four mass
blocks connected with sail ends to extend due to
centrifugal force.
Both configurations of the ADM use a similar
technique to deploy their stowed sails. Based on this
and the previously obtained result, four critical
components were chosen. These are 1. Sail, 2. Rotator
3. Connector.
The first element chosen encompasses the sail
configuration and the characteristics of its membrane.
The influence of this element of the design requires
selection of the configuration of sail. A square shaped
configuration was chosen due to providing the best drag
to mass ratio as well as having a broad-space heritage.
The chosen sail configuration is a single unit with
intrinsic divisions into 4 triangle quadrants. This
configuration is perfect for deployment using angular
momentum as the individual mass tips could uniformly
spread out the sail and stay at maximum distance on its
diagonal. The average size of sail to be used would stay
7m2. The proposed material of the sail is a metalized
polymer similar to the one used for the sail of the
CANX-7 mission. The polymer proposed to be used is
a 12.7 micrometer Kapton film with a 300-Angstrom
Aluminum coating capable of handling greater than
200° temperatures. This makes the mass/m2 equal to ~7
for cubesat configuration and ~4 for sat configuration.
The aluminum coating allows achieving high
equipotential space-craft structures and avoids
excessive charging of membrane. The material
selection was in accordance with COTS philosophy and
objective to use tested components. Secondly, since no
boom deployment mechanism will be used, the absence
of boom configuration would otherwise cause the
deployed sail to fold in on itself as the masses loose
tension. The spacecraft will not be kept in a constant
revolving state as the increase in mass would mean
increased power usage to keep a controlled spin rate; it
increases communication hindrances. To solve this, a
photopolymer in the form of epoxies or nitrile rubber
would be used in the sail along its diagonals (fig. 2).
Photopolymers harden under sunlight undergoing a
process called curing [10]. The material selection again
follows COTS philosophy and a preference to those
already tested in space would be given. To be noted,
this is only for cubesat configuration.
Figure 2: Photopolymer application at
diagonals of sail
The second element of importance is the Rotator. It
derives aspects of stowage from IKAROS probe
design. The primary functions of the Rotator include
stowage and deployment. The Rotator element is a
cylindrical probe of diameter and length not exceeding
20 cm in both configurations. The sail is folded in a
rolled 4 petal fashion similar to IKAROS. The
IKAROS deployment mechanism utilizes the 4
stoppers which aid in deployment of the rolled petals
using relativistic rotation mechanism (motor drive) and
eventually are released to allow the expansion of then
fully deployed petals into a square sail. The tip masses
attached to sail ends are also stowed in Rotator and are
the first ones to deploy on rotation aiding in the
deployment process by the tension created due to
experiencing centrifugal force. Each tip mass is 50g in
mass. Angular momentum to deploy the sail is
generated by a reaction wheel. Adhering to COTS and
space tested component requirements, the reaction
wheel suggested is CubeSpace™’s CubeWheelⓇ
Small+. It is the only rotating module and is further
connected to the host space-craft by a Connector
module.
The Connecter module is a passive module, not
present in cubesat configuration, connecting the
Rotator to the host satellite. It is fixed to the satellite on
one end and extends a cylindrical rod until the end of
the Rotator. The Rotator rotates around this rod of
connector as the axis. The Connecter furthermore
houses a housekeeping computer controlling the entire
module primarily providing needed power control. To
dump excess momentum and avoid tumbling in the
satellite caused by this rotation, the connector consists
of magnetorquer rods. Components chosen based on

COTS philosophy and space tested systems are NCTR-
M002 Magnetorquer Rod. This isn't present in cubesat
configuration as the Rotator is directly attached to the
satellite if it is capable of holding together during the
deployment phase. This is done to keep the mass and
power usage low. Figure 3 shows the design of
Connector and Rotator.
Figure 3: Connector, Rotator, Rotator Rod
4.2 Deployment Mechanism
Deployment process can be initiated from ground or
host spacecraft. Once initiated, the power supply to the
reaction wheel is regulated and increased to have the
Rotator rotate at 5 rpm over a span of 6 seconds. As the
rotation speed increases gradually, at 2 rpm the tip
masses are released which are clutched mechanically.
This creates a centrifugal force experienced by the tip
masses which extends the sail-petals gradually
connected to the Rotator by tethers. The spin rate is now
increased to 25 rpm as the petals extend out while
stoppers hold the membrane through the relative
rotation mechanism. The spin rate gradually decreases
as the petals extend half their size to ~15 rpm as no new
torque is being produced. The spin rate on full
extension of petals is ~3 to 4 rpm and on full
extensions, the stoppers are released. As the stoppers
“fall down”, the final stage of deployment begins and
the sail starts acquiring its square shape from petals. By
this point, the rpm is low at ~2. In cubesat
configuration, this rate is kept for an hour under
sunlight post deployment for the sail to harden at its
diagonals after which the dumping devices are used to
dump the rotation. In the other configuration, the
Rotator is kept spinning at the rate as the Connector’s
dumping magnetorquer dumps excess rotation and
draws power from the satellite. Figure 4 represents the
deployment sequence.
Figure 4: Deployment Sequence
4.3 Design Overview
To fulfill its initial objectives, the design should
provide the said results within the cost and mass
constraints minimizing both of them. Table 1 gives an
estimation of cost and mass for both configurations,
also justifying the division of design into two
configurations visible by the considerable mass and

cost difference mentioned in the table apart from the
size difference described in Conceptual Design. The
components are COTS and space tested on multiple
missions previously thus adhering to initially said
philosophies.
Table 1
Components and Cost
Component
Mass
Estimated Cost (USD)
Cubesat
Config
Chassis
700g
2500
Reaction wheel
90g
5950
Sail
~50g
500
Dumping rod
30g
1200
Electronics
1500g
700
NET: ~ 1100g
NET: ~10k USD
Sat Config
Chassis
1000g
~3100
Reaction wheel
90g
5950
Sail
~28g
250
Dumping rod
30g
1200
Electronics
200g
800
NET: ~ 1400g
NET: ~11k USD
5. Conclusions
This research offers an alternative deployment
mechanism conceptualized in the form of above-
mentioned design which is clearly plausible in theory.
More testing of the mechanism in controlled
environments along with prototyping and testing the
design should be the next steps in developing a full-
fledged design. The mechanism and sequence given

could be schematized and tested on the prototype.
More research is needed towards various deorbiting
mechanisms to avoid increasing the already growing
space debris problems.
6. Nomenclature
𝐹𝑑 = drag force,
𝜌 = density of the medium,
𝐴 = cross-sectional area i.e., area perpendicular to
atmospheric flow vector,
𝐶𝑑 = the coefficient of drag
𝑣 = orbital velocity of craft
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