CubeSat Deorbit Devices
Final Report
Old Dominion University
Project Design & Management II
April 26, 2011
Dr. Robert Ash
Lindsey Andrews
Jake Tynis
Joshua Laub
Abstract:
The integration of electronic devices and subsystems within a satellite bus is not a trivial issue in the design of a spacecraft. The benefit of this research is immediately apparent, expanding experiment opportunities using low-cost access to space utilizing the CubeSat platform. However, once orbit has been achieved, a CubeSat's lifetime is finite and can be on the order of decades. This research has shown that an effective deorbit device can be deployed to shorten the lifetime of a CubeSat. This technology is becoming more valuable as the total number of CubeSats in orbit increases. The future safety of the earth orbit environment with regards to orbital debris depends on the ability to safely deorbit nanosatellites.
Table of Contents
List of Figures
Introduction
I. Literature Review
II. Rationale
III. Project Objective
IV. Benefits
Proposed Approach
Finalized Approach
I.Materials
II.Testing
III.Valve
IV.Software
A.Satellite Tool Kit
B.CubeSat Plug and Play Development Kit
C.LabVIEW
D.AVR Studio
V.Machine Shop Work
VI.Vacuum Chamber
VIII.Digital Microcontroller
Organization
Final Deployment Testing
Cost Consideration
Future Recommendations and Research
Summary
Appendix
I. Works Cited
II. Secondary Material
III. Microcontroller Source Code
List of Figures
Figure 1: Six CubeSats and their respective P-POD launcher (1)
Figure 2: COTS CubeSat Structure (2)
Figure 3: Predicted lifetime of a 1U CubeSat without deorbit module (7)
Figure 4: Predicted lifetime of a 1U CubeSat with deorbit module (7)
Figure 5: Deorbit Pillow Construction (7)
Figure 6: Specifications for CubeSat adhesive
Figure 7: Corner Folding Method (7)
Figure 8: Test Photo
Figure 9: Valve and Manifold Assembly
Figure 10: Machine Shop Drawing
Figure 11: 1U CubeSat Shape
Figure 12: Volume Requirement
Figure 13: Bell jar (with metal bell jar guard) for simulating space conditions
Figure 14: Plot of semi-minor axis as a function of internal pressure
Figure 15: Geometry of blister test problem
Figure 16: Values for semi-minor axis and burst thickness
Figure 17: Container lid for deployment testing, showing cut-outs for stress concentration
Figure 18: ATMega32 Pinout
Figure 19: Lab electronics equipment
Figure 20: Primary and secondary failure locations of container lid under inflation
Figure 21: Prototype Budget
Figure 22: Space Qualified Budget
Introduction
I. Literature Review
Low cost access to space has been a driving force since the early days of spaceflight. Over the decades, advances in technology have allowed space structures and avionics to decrease in size. The CubeSat architecture takes advantage of both of these advances.
The CubeSat satellite standard resulted froma collaborative project between California Polytechnic and State University, San Luis Obisbo and Stanford University. A standard1U CubeSat is a 10 centimeter cube with a maximum total mass of one kilogram or less. The CubeSat design is based on the premise that low-cost access to space can be achieved when these systems are deployed as a secondary payload to a host launch system. In other words, if extra mass is available after the primary spacecraft and associated hardware are designed for a particular launch vehicle, CubeSats can be deployed from the same launch vehicle stack for a very small additional cost. Unfortunately, many of the CubeSat these “satellites of opportunity” have been deployed at orbital altitudes on the order of 900 km, where the CubeSat satellite orbital lifetime can be hundreds of years.
CubeSats are deployed using the Poly-Picosatellite Orbital Deployer, P-POD, as shown in Figure 1: Six CubeSats and their respective P-POD launcher (1), also developed by California Polytechnic and State University (1).
Figure 1: Six CubeSats and their respective P-POD launcher (1)
As stated previously, the standard size for a CubeSat is a 10 cm cube (referred to as 1U). This can actually be increased to a maximum size of 10 cm by 10 cm by 30 cm and 3 kilograms (3U). Obviously this is a larger, more expensive endeavor than a standard 1U CubeSat; however it is an option.
The design challenges of CubeSats are very apparent: small size. The small size and mass requirements seriously constrict the payload. The only volume and mass remaining for the experiment is what is left after essential systems have been integrated in the bus. There must be an electrical power system, telemetry, data handling, thermal control, and various other systems that must be incorporated regardless of the actual experiment.
Companies now specialize in the development and sale ofcommercial off the shelf (COTS) CubeSat subassemblies, as shown in Figure 2: COTS CubeSat Structure (2).These kits greatly aide in reducing the overall payload development time and enable precise budgeting of many spacecraft elements. The main idea behind CubeSats is to keep cost and development time to a minimum.
Figure 2: COTS CubeSat Structure (2)
The launching of artificial satellites into earth orbit has produced some unintended consequences. Every satellite has a finite useful lifetime; at some point they will no longer function and thus become debris. Furthermore, a variety of upper launch stages and mating systems often achieve stable long-life orbits even when they have no function other than placing the spacecraft into orbit. As the total number of orbiting satellites andassociated debris increases, attention must be focused on minimizing their orbital lifetime. There are currently over two million kilograms of space debris in orbit around the earth (3). Orbital debris can be divided into three distinct groups :(1) accidental or intentional break-ups; (2)intentional release of objects from launch vehicles and spacecraft during deployment; and (3) in-orbit collision-derived creation of space debris (4). These three separate categories may result in objects that have lifetimes greater than 1000 years.
The United Nations Office for Outer Space Affairs has prescribed a series of guidelines which are designed to mitigate orbital debris. The guidelines are as follows:
- Limit debris released during normal operations
- Minimize the potential for break-ups during operational phases
- Limit the probability of accidental collision in orbit
- Avoid intentional destruction and other harmful activities
- Minimize potential for post-mission break-ups resulting from stored energy
- Limit the long-term presence of spacecraft and launch vehicle orbital stages in the low-Earth orbit (LEO) region after the end of their lifetime (4)
Guideline six requires that launch stages, their associated hardware and payloads must be designed to result in a timely return to earth at the conclusion of their mission. Additionally, the Inter-Agency Space Debris Coordination Committee (IADC) along with NASA and the International Standards Organization (ISO) put a limit on orbital lifetimes for Low Earth Orbit (LEO) of 25 years(5).
II. Rationale
The goal of this design project was to develop an integrated CubeSat deorbitdevice that when deployed could cause CubeSats to deorbit from a 900 km altitude in less than 25 years. Designs for deorbit mechanisms range from long tethers to inflatable assemblies. The current level of deorbit mechanism research provides many new areas to explore.
The method chosen to best meet the requirements for a CubeSat deorbit device is one that increases surface area, producing drag. This design relies on creating a drag force to slow the orbital velocity of the spacecraft(3). It can be shown that as the orbital velocity decreases, the orbital altitude must also decrease. The drag force is a function of the atmospheric density at the specific altitude. Atmospheric density varies between night and day, and is affected by solar activity. However, in general, atmospheric density decreases exponentially from the earth’s surface outward; therefore drag is higher at lower altitudes.
In spacecraft design, ballistic coefficient can be used to describe the drag efficiency. Ballistic coefficient is defined as where M is the mass of the orbiting object, CD is the drag coefficient, and A is the frontal area (6). In this case, a lower ballistic coefficient will have a larger surface area and be a better deorbit mechanism. A 1U CubeSat, with a mass of 1 kg, a cross sectional area of 0.01 m2, and a CD of 2, will have a ballistic coefficient of 50.
The method for increasing the frontal surface area of the spacecraft requires some sort of deployable mechanism. The two main ways to increase the surface area of a spacecraft are to deploy a rigid structure/array or an inflated device. Both of these methods require some sort of activation technique and movement.
The deployable array technique would be similar to deploying additional solar panels. The sides of the spacecraft could fold outward to increase the total surface area. This technique required a significant level of sophistication and power, and was discarded.
The second method for increasing surface area was to inflate a generic volume. From a drag perspective, the shape was irrelevant as long as the frontal surface area for producing the required spacecraft lifetime was achieved. An inflatable must use some sort of inflation gas. The thin-walled inflatable membrane and associated inflation gas and hardware can only occupy a small volume in order to provide sufficient volume for the actual payload. Consequently, the deorbit system, including the inflation gas and activation system must be capable of being packed into avery small volume. Material selection for the thin walled inflatablewas also of importance since it must remain in space for approximately 10 years. Puncture resistance was a consideration, regardless of whether micrometeoroid impacts are a concern or not (6).
A possible modification to the inflatable systemwas considered utilizing a conductive ring inside the membrane material. The ring would allow a current to flow, which would interact with the Earth’s magnetic field. This interaction could be beneficial if the conductive ring was perpendicular to the Earth’s magnetic field lines, since then the ring would experience a torque, which could be used to orient the CubeSat. Alternatively, if the ring wasoriented parallel to the magnetic field lines, the application of a suitable DC voltage could produce a current (clockwise versus counterclockwise), and the interaction could cause the ring to expand. This expansion could be useful. For example, if the inflatable device experienced a leak, the electromagnetic expansion could helpkeep the material from collapsing on itself, making it easier for the gas to inflate the structure.
Although a conductive ring was considered useful, the added complexity was a concern. One would need to ensure that dissipated heat from the ring did not degrade the adhesive in the inflatable structure. The folding and packaging of the material would need to be slightly altered. Also, the CubeSat would need to store more post-mission power, in order to supply the current to the ring. Lastly, and most importantly, a detailed numerical model would be necessary to predict the magnetic field of the Earth at the exact position of the CubeSat. The purpose of this computer program would be to ensure that the magnetic forces were directedappropriately ( according to the right-hand rule). Due to these complexities, a conductive ring was eliminated from further consideration.
III. Project Objective
The objective of this design project was to construct a CubeSat deorbit device. Using applicable research sources, a best practice design for an inflatable deorbit device was to be constructed and demonstrated as a prototype. The device was to be part of the payload for a 1U CubeSat, meeting all of the standards for that CubeSat type.
IV.Benefits
The benefits of constructing this demonstration device was a proof of concept. The first concept to be proved was that this payload can meet the constraints of a 1U CubeSat payload (size, mass, etc.). The second benefit was the development of a low cost, reliable method to deorbit a CubeSat. The removal of high altitude CubeSats will reduce a significant future source of space debris.
Proposed Approach
The approach chosen was to design an inflatable balloon or pillow shape. This design was less complicated than a fixed folding array design. Depicted in Figure 3: Predicted lifetime of a 1U CubeSat without deorbitmodule (7). It can be seen in the figure that the predicted lifetime is well in excess of 400 years. Figure 4 is the predicted lifetime of a 1U CubeSat with a ballistic coefficient produced by increasing the frontal area to approximately 0.6 m2 . The increased surface arearesulted in an effective deorbit device. The deorbit device successfully removed the satellite from orbit within the prescribed 25 year life span.
Figure 3: Predicted lifetime of a 1U CubeSat without deorbit module (7)
Figure 4: Predicted lifetime of a 1U CubeSat with deorbit module (7)
The actual shape of the inflatable structure is dictated by the complexity of constructing the volume. It is for this reason that a pillow shape, as shown in Figure 5, will be used. The pillow design need only be sealed along the edges, and thus lends itself to a simpler design. Material selection of the pillow is critical. In the harsh low earth orbit regime, radiation and atomic oxygen are the primary concerns (7). The material chosen to survive this environment is a polyimide film known as Kapton®. Kapton is a product of the DuPont Company and has a proven track record for long term space applications (8).
The inflation of this balloon must be done with a reliable gas source. There are two main options in this area. The first is a cool gas generator, and the second is a stored refrigerant. Several companies specialize in developing space qualified cool gas generators for various applications. One company specifically, Bradford Engineering, has developed a 2 gram cool gas generator. This is a perfectly viable method to inflate a small deorbit mechanism for a CubeSat (9). For the refrigerant option, a simple pressure vessel containing the fluid would be sufficient (7).
Figure 5: Deorbit Pillow Construction (7)
The actual triggering of the deorbit device can be accomplished either by an onboard timer or with ground based signaling. The actual activation will be processed through the main bus of the CubeSat and should require a minimal fraction of the total CubeSat’s available power (~5%). The housing for the deorbit device will be constructed of the 6061-T6 aluminum alloy, per NASA and California Polytechnic guidelines. This alloy is common to the entire CubeSat and P-Pod launcher, and offers the best density option for design integration and weight minimization.
Finalized Approach
It was determined that the prototype deorbiting system was to be an inflatable device. Using rigid structures to increase the surface area would be impractical due to the added mass of the rigid structure, actuators, gears, motors etc. In addition, inflation can occur as a passive action. Once the valve to the pressurant gas is opened no further electrical or mechanical input is necessary. With the rigid structures enough battery reserve must be maintained to run the electric motors/actuators to drive the gears and tracks. Electrically intensive experiments may not provide enough end-of-life power margins to accommodate that. There are also fewer moving parts with the inflatable. For example launch loads may accidentally misalign gears and tracks or electric motors may seize with the rigid structure whereas the lid and valve are the only moving inflatable parts and are inherently simple.
I.Materials
As mentioned previously, the primary materials required for the successful demonstration of the deorbiting technology include: aluminum 6061-T6, Elastosil S36 adhesive, Suva 236fa refrigerant, and Kapton polyimide film. During the course of designing and fabricating the prototype, as with all projects, substitutions and compromises had to be made. These compromises were approached from the standpoint that the purpose of the project was to demonstrate our ability to communicate with the CubeSat bus and trigger the deployment of an accurately sized and operational deorbiting device.
While NASA regulations require that structural components for CubeSats be manufactured from certain alloys of aluminum, the current prototype would not actually be utilized on a launch vehicle. With that being said the decision was made to work with the machine shop and utilize the materials they had on hand. The design constraint was that the deorbit device enclosure must still fit within the size restrictions as originally outlined. The added benefit of this decision was that the container was fabricated from stock, effectively eliminating the cost of procuring aluminum which was originally estimated to be $15-$20. Costs for aluminum were subsequently verified upon contact with several suppliers throughout the greater Hampton Roads area as well as nationwide. Prices ranged from as low as $4 per sheet for a couple square feet to as much as $146 for 48 ft2. Substituting stock material for aluminum 6061-T6 was deemed acceptable as it will not impact the overall success or failure of the prototype demonstration.
For demonstration purposes, the inflatable device must be activated similarly to how it would be when actually deployed on the satellite. Developing the prototype as an integrated whole and also incrementally in stages requires many test inflations. Utilizing Suva 236fa for each test firing was determined to be impractical; each time a test was to be performed the small cartridge would have to be charged with new refrigerant and sealed. The lab in which testing and integration is carried out provided an option to eliminate the cylinder recharging from the routine. A metered air supply from a large reservoir is fed to the lab via air line, which has a built in pressure gauge. Air is available in pressures ranging from just barely above atmospheric to as much as 100 psi. This pressure range is more than viable for the inflatable. On orbit, the Suva would be providing a pressure of 84.7 psi for inflation into a vacuum. To compensate for sea level the pressure on the air supply can be metered to 99.4 psi mimicking the Suva and also compensating for ambient conditions. Using the lab air supply was a conclusion based on both necessity and practicality: necessity because attempts to contact a Suva 236fa supplier were unsuccessful, and practical for the reasons outlined above. Success of the prototype demonstration will not be impacted by this substitution and the results will actually be improved. Improvement will come from the fact that the pressure supplied to the inflatable can be more readily controlled with the air supply as opposed to a small gas cylinder whose pressure is sensitive to temperature fluctuations. Also, the effects of under pressurization and over pressurization can be easily studied by simply increasing the flow from the air supply. A benefit in project cost is seen in the elimination of purchasing refrigerant, obtaining a suitable gas cylinder, and EPA regulations.