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Verification of MEMS fabrication process for the application of MEMS solid propellant thruster arrays in space through launch and on-orbit environment tests
Author’s Accepted Manuscript Verification of MEMS Fabrication Process for the Application of MEMS Solid Propellant Thruster Arrays in Space through Launch and On-orbit Environment Tests Hyun-Ung Oh, Tae-Gyu Kim, Sung-Hyeon Han, Jongkwang Lee www.elsevier.com/locate/actaastro
PII: DOI: Reference:
S0094-5765(16)30629-4 http://dx.doi.org/10.1016/j.actaastro.2016.11.013 AA6077
To appear in: Acta Astronautica Received date: 3 July 2016 Accepted date: 8 November 2016 Cite this article as: Hyun-Ung Oh, Tae-Gyu Kim, Sung-Hyeon Han and Jongkwang Lee, Verification of MEMS Fabrication Process for the Application of MEMS Solid Propellant Thruster Arrays in Space through Launch and Onorbit Environment Tests, Acta Astronautica, http://dx.doi.org/10.1016/j.actaastro.2016.11.013 This is a PDF file of an unedited manuscript that has been accepted for publication. As a service to our customers we are providing this early version of the manuscript. The manuscript will undergo copyediting, typesetting, and review of the resulting galley proof before it is published in its final citable form. Please note that during the production process errors may be discovered which could affect the content, and all legal disclaimers that apply to the journal pertain.
Verification of MEMS Fabrication Process for the Application of MEMS Solid Propellant Thruster Arrays in Space through Launch and On-orbit Environment Tests
Hyun-Ung Oh1, Tae-Gyu Kim1, Sung-Hyeon Han1, and Jongkwang Lee2*
1
Department of Aerospace Engineering, Chosun University, 375 Seosuk-dong, Dong-gu,
Gwangju, 501-759, Republic of Korea 2
Department of Mechanical Engineering, Hanbat National University, 125 Dongseodaero,
Yuseong-gu, Daejeon, 305-719, Republic of Korea
*
Corresponding
author,
[email protected]
ABSTRACT
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One of the most significant barriers encountered to the space application of MEMS technology is its lack of reliability and flight heritage in space environments. In this study a MEMS solid propellant thruster array was selected for the verification test of MEMS technology in space. The function and performance of MEMS solid thruster have been previously verified by laboratory-level research in universities. To ensure the successful operation of the MEMS thruster module before flight demonstration on-orbit, launch and onorbit environment tests were performed at the qualification level. In the launch test, sine burst, and random vibration loads were applied to the MEMS thruster module. The thermal vacuum tests were carried out for the on-orbit environment test. As a result of the launch vibration test and on-orbit environment test, the variations of the characteristics were less than 0.7%, and all the functional requirements were successfully verified after the vibration tests. The tests successfully verified the manufacturing process because the thruster module showed stable normal function before the ignition. The test result outputs will be helpful in establishing MEMS fabrication guidelines for space applications.
Keywords: MEMS Solid thrust, Cube satellite, Launch vibration test, Thermal vacuum test
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1. Introduction
The mass, volume, and power consumption of an on-board mission payload is a significant contributor to the overall increase in the cost of spacecraft systems. Microelectromechanical systems (MEMS) technology offer several advantages for space applications owing to the drastic reduction in size, weight, and power consumption, made possible by the miniaturization of components and systems [1]. Recently, MEMS is being considered for use in micro-sensors, micro-propulsion systems, and RF switches in space systems for which such devices are of specific interest [2–10]. MEMS systems increase the functionality of satellite systems with limited volume while simultaneously decreasing spacecraft weight. However, to realize the ultimate goal of sending completely micro-fabricated and integrated MEMS-based devices into space, the long-term reliability of MEMS technology needs to be assessed by actual experiments in space, and this will require frequent launches and on-orbit verifications in the near future. Cube satellites (CubeSats) are useful tools for building the space heritage of MEMS technology in on-orbit through orbital experiments and verification tests of new technologies. These satellites are cube-shaped, pico-class miniaturized satellites, and are much smaller than commercial satellites. They usually have a volume of 10 cm3, and a 1
mass of less than 1.33 kg for a standard size of 1U [11]. Recently, CubeSats have been widely advocated by organizations and universities around the world to achieve increasingly
complex
scientific,
surveillance,
interplanetary,
and
technology
demonstration missions. CubeSats capable of orbit control can be used for a wide range of scientific missions; their functionality in an extremely small package promises to yield numerous advanced technologies required for achieving challenging missionrelated functions. In order to realize these missions, new types of micro-thrusters [12– 15] that offer a wide range of impulse bits with much lower total mass are required. The development of micro-thrusters has been actively researched in the recent years, and the performance of various types of thrusters has been investigated. Santoni et al. [12] developed a MEMS-based cold gas micro-propulsion system for attitude control to be tested on board the Ursa Maior CubeSat within the framework of the QB50 projects. The electrolysis propulsion systems [13] investigated by Cornell University are well suited for CubeSat-scale missions because of their simplicity, flexibility in terms of electrical power consumption, and dense propellant storage. These systems provide an estimated 1 km/s of V, which is a significant improvement in the capabilities of CubeSats. A vacuum arc thruster for CubeSat propulsion is also scheduled for an inspace test on board the UIUC CubeSat [14]. Lee et al. [15] proposed an improved
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MEMS-based solid propellant thruster array with a micro-membrane igniter. They improved the stability of the micro-igniter by using a glass membrane tens of microns thick. They also developed a process to fabricate the micro-igniter on a photosensitive glass wafer instead of on a dielectric membrane. STEP Cube Lab (Cube Laboratory for Space Technology Experimental Project) is the first pico-class satellite being developed at the SSTL (Space Technology Synthesis Laboratory) of Chosun University, and it is scheduled to be launched in 2016 [16]. The main objective of this mission is to perform on-orbit verification of results of spacerelated research conducted at domestic universities. The MEMS-based solid propellant thruster array of Lee et al. [15] is one of the main payloads to be verified through the STEP Cube mission. The main purpose of on-orbit verification of MEMS-based solid propellant thruster arrays is to verify the MEMS fabrication process for use in space environment rather than to verify the capability of the attitude determination by the MEMS propulsion system. This is because the MEMS technology is very attractive for space applications as it offers great benefits such as reduction in size, mass, and power consumption through miniaturization of components and systems. However, to realize the ultimate goal of sending completely micro-fabricated and integrated MEMS systems into space, the long-term reliability of this technology needs to be assessed through
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experiments in space, and this will require frequent launches in the near future. The MEMS fabrication processes of a solid thruster array such as the sputtering of Pt/Ti heater, diffusion bonding of photosensitive glass, ultraviolet (UV) bonding, solid propellant filling under vacuum condition and electrode design approach will be verified through this program. The successful activation of nine solid thrusters at specified time intervals during life time is used as the success criterion to judge whether the aforementioned MEMS fabrication processes are applicable for space applications. In this paper, we propose a mounting method that employs flexure-like brackets to support the MEMS thruster module, which consists of a MEMS solid thruster and its control board. This method makes it possible to reduce the launch loads transmitted to the thruster module under launch environments. It also allows the thruster module to be easily mounted on or dismounted from the fully integrated satellite structure without disassembling any of the satellite parts during the test and transportation to the launch site. In addition, we propose an electrode pad design combined with spring-loaded pogo-pins instead of soldering of electric wires, which is generally applied in the conventional MEMS fabrication process, to achieve reliable electrical contact between the thruster and its control board under launch and on-orbit environments. In the study, to guarantee successful operation of MEMS thruster on-orbit and to verify the
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fabrication process including the modified design for space usage, space qualification tests were conducted before flight demonstration. The results obtained from the harsh launch vibration test and thermal vacuum environment test will be helpful for establishing the MEMS fabrication guidelines for space applications.
2. Overview of STEP Cube Lab
Figure 1 shows the mechanical configuration of the STEP Cube Lab; its design is based on a 1U standard CubeSat. The main objectives of the STEP Cube Lab mission [16] are to identify core space technologies studied in domestic industries or universities, and to verify them in orbit. Its primary objective is to perform on-orbit verifications of fundamental space technologies for future space missions. The payloads to be verified in this mission are a variable emittance radiator, a phase change material (PCM), a MEMS-based solid propellant thruster [15], a concentrating photovoltaic (CPV) power system [17], and a novel non-explosive holding and release mechanism triggered by nichrome burn wire heating [18]. The MEMS-based solid propellant thruster located at the bottom of the CubeSat is also one of the main payloads to be verified through the STEP Cube Lab. mission.
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3. MEMS Solid Propellant Thruster Array
3.1 Design Modification of MEMS Solid Thruster for Space Usage Figure 2 (a) shows a cross-sectional view of the 3×3 MEMS solid propellant thruster array. It consists of four layers of photosensitive glass, a micro-nozzle layer, an intermediary layer, a micro-igniter, a seal layer, and the solid propellant. The micronozzle is located on the first layer, and the second layer is the intermediary layer. The third layer contains the micro-igniter, with a glass membrane tens of microns thick, and the solid propellant. The fourth layer acts as the seal layer. The operation principle is as follows: when the solid propellant attains its ignition temperature achieved by the micro-igniter, the glass membrane is subsequently broken by the high-energy combustion gases of the solid propellant. Thrust is generated through the micro-nozzle. The specifications of the MEMS thruster array are summarized in Table 1. Figure 2 (b) shows the conventional glass-ceramic membrane igniter and Fig. 2 (c) shows the improved glass membrane igniter. In the current design, in order to improve the ignition performance by preventing ignition failures caused by incomplete combustion due to delay in ignition time, the micro-igniter is installed under the glass membrane. Therefore, the solid propellant can be more efficiently ignited because the micro-igniter in the modified design is in direct contact with the solid propellant unlike in the conventional MEMS thruster, where the propellant is heated-up indirectly by an igniter installed on the glass membrane. A numerical simulation of the modified MEMS igniter design for increasing the
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success rate in orbit was carried out to predict its ignition characteristics using the commercial code, Fluent. The temperature change of the propellant as a function of the time was monitored at 15 V, the required voltage for the ignition of the solid thruster. Figure 3 shows a comparison of the cross-sectional temperature contours between the proposed direct-heating igniter and the conventional indirect-heating igniter at 20ms. The temperature increase rate at the top surface of the propellant was faster for the proposed igniter than for the conventional igniter owing to the position of the heater. The ignition delay and ignition energy of the proposed igniter as a function of the input voltage are presented in Fig. 4. The ignition characteristics depended on the input voltage. The ignition delay and the ignition energy decreased exponentially as the voltage increased. The numerical results showed an ignition delay of 16.8 ms, and the ignition energy of 24.9 mJ at 15 V; these values are about 7.4% lower than the corresponding values for the conventional igniter. In this study, we also proposed a modification of the design of the electrical interface of the MEMS solid thruster to guarantee reliable electrical contact between the thruster and its controller. The electrical interface on the thruster was modified such that it was directly connected to the pogo-pins on the thruster control. The details are described in Section 3.3.
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3.2 Fabrication Process of the MEMS Solid Propellant Thruster The fabrication process of MEMS thruster was divided into three steps: the fabrication of the micro-nozzle, intermediary, and seal layer; the fabrication of the micro-igniter; and the final assembly process, including the loading of the propellant. The fabrication process of the micro-igniter is shown in Fig. 5. In the first step, the micro-nozzle layer, intermediary layer, and seal layer were fabricated by anisotropic etching of photosensitive glass. The nozzle layer was fabricated by anisotropic etching on one side to obtain the divergence micro-nozzle. The average diameter of nine nozzle throats was 416 μm. The average area ratio of the micro-nozzles was 1.85. The thickness of the micro-nozzle layer, intermediary layer, and seal layer were 0.9 mm, 0.3 mm, and 0.9 mm, respectively. Next, the micro-igniter comprising a heater, a membrane, and a propellant chamber was fabricated. Platinum (Pt) was adopted as the material of the heater because it exhibits high stability at high temperatures and is resistant to oxidation and corrosion. A 40 um thick glass membrane was selected from the viewpoint of structural stability. Glass membranes have lower thermal conductivity and lower fabrication costs than conventional dielectric membranes. The micro-igniter was fabricated by anisotropic wet etching, Pt/Ti deposition and patterning, chemical-mechanical polishing (CMP), and UV bonding process. The glass was then selectively exposed to UV light of wavelength of approximately 310 nm and energy density 2.5 J/cm2. Then, the exposed glass wafer was inserted into a programmable furnace for heat treatment. The glass was heated to
8
500°C at a ramp rate of 2°C /min, and the temperature was maintained constant for 1h. Then, it was heated again to 585°C with a ramp rate of 1°C /min and maintained at this temperature for 1 h. Finally, the glass was cooled at a cooling rat e of −3°C /min. During this heat treatment, the properties of the UV-exposed glass changed because of re-crystallization. After the heat treatment, the wafer was soaked in a diluted 10% hydrofluoric (HF) acid solution. The general etching ratio of glass-ceramic to vitreous glass is 20:1. Because the unexposed area was partially etched, the surface of the wafer was roughened. Subsequently, the wafer was chemically and mechanically polished in order to make its surface flat. The membrane layer was fabricated by patterning the heater and electrodes onto the glass wafer. First, a photoresist was spin-coated onto the glass and patterned by photolithography to form the heater and electrodes. A layer of 200 Å thick Ti layer was sputter deposited to improve Pt adhesion to the glass. A 2000Å thick Pt layer was sputter deposited on the Ti layer. The photoresist on the glass and on the Pt/Ti layer were removed by applying acetone for 2 min. After this process, only the Pt/Ti pattern of the micro-igniter remained. After each layer of the igniter had been fabricated, the membrane layer and the chamber layer were integrated by a UV bonding process. The UV glue was spin-coated onto the membrane layer. The membrane layer and the chamber layer were then aligned, and exposed to UV light of wavelength 310 nm. The fabricated micro-nozzle layer and intermediary layer were polished and cleaned in piranha solution for 10 min. Two layers were pressed together at 500°C by
9
applying a pressure of 2 kPa for 12 h with a ramp rate of 2°C/min and a cooling rate of about −3°C/min. Lead styphnate was selected as the solid propellant, considering its ease of loading and low power consumption for ignition. The solid propellant was dissolved in a solution that contained the solvent and the binder. The solution was put in the chamber layer and dried to remove the solvent. After the filling process, all layers were assembled by UV bonding because the propellant is vulnerable to heat exposure. Figure 6 shows each fabricated layer of the MEMS solid propellant thruster and the fabricated micro-igniter.
3.3 Design of MEMS Solid Thruster Assembly for Space Applications In this study, to achieve a reliable electrical contact between the MEMS thruster and its control board in launch vibration and thermal vacuum environments, an electrode pad design combined with spring-loaded pins instead of the conventional soldering electric wire is proposed. Figure 7 shows an exploded schematic view of the MEMS thruster module and the design of the electrical interface between the thruster and its control board for controlling the activation of the propellants according to the ignition commands on-orbit. The figure also shows the modified electrical interface on the backside of the solid thruster. The conventional design, employing electric wire soldering [15] generally used in the MEMS fabrication process is shown in Fig. 8. However, this approach might not be feasible for space applications considering the harsh launch vibration loads and thermal loads induced by temperature variations on-
10
orbit. Therefore, the conventional design was slightly modified for space applications as follows. To achieve reliable electrical contact and survivability, we proposed electrode pads design combined with spring-loaded pogo-pins. As shown in the exploded schematic view in Fig. 7, the electrical interface on the thruster was directly connected to the pogo-pins on the thruster control board and fixed by a plastic cover made of Delrin® with space heritage. The function of the thruster module was tested by the external command to the thruster control board. The ignition test performed under ambient conditions at the MEMS thruster assembly level showed successful ignition of all nine thrusters. Figure 9 shows the images captured by a high-speed camera, showing the successful ignition of the thruster activated by the control board; the images indicate that the ignition was successful without any malfunction under ambient conditions. In addition, to ensure successful on-orbit verification of the MEMS thruster array, structural design effort at the satellite level are performed to minimize the launch loads transmitted to the thruster under launch environments. This was achieved by mounting the MEMS thruster module on flexure-like brackets made of Delrin® plastic material, as shown in Fig. 1. This design also makes it easier to access and replace the MEMS thruster from outside without de-mating the other parts of the CubeSat because frequent replacements of the thruster are required during the verification test. For example, a dummy thruster without solid propellant will be used to prevent ignition during a full function test of the system level thermal vacuum test. This will also be
11
replaced by a flight model of the thruster with solid propellant before the launch campaign.
4. Space Environmental Qualification Test
One of the most significant barriers encountered to the space usage of MEMS technology is its lack of reliability and flight heritage in space environments. The function and performance of MEMS solid thruster have been previously verified by laboratory-level research in universities, but their design effectiveness has never been qualified in outer space or on-ground simulated space environments. Therefore, it is unclear whether MEMS thruster will operate without any malfunction in an on-orbit environment. In this study, to guarantee the structural safety and successful operation of MEMS thrusters under severe launch and on-orbit environments, we performed a vibration test and thermal vacuum test under qualification level. The main objectives of the qualification tests were to verity the design and fabrication process of the MEMS solid thruster described in the Section 3.2 for judging whether the technology can be used in space applications before actual flight demonstration. The modified design of the electrical interface between the thruster and control board was also verified in the qualification tests. In the vibration test, the effectiveness of the structural design employing a flexure-like bracket design was also investigated.
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4.1 Launch Vibration Test Figure 10 shows the vibration test set-up for the qualification test of the MEMS thruster module. In the test, sine burst, and random vibration loads were applied to the MEMS thruster module. The qualification test level was based on the QB50 test specifications summarized in Table 2 [19]. An accelerometer sensor, which is used to apply input test loads was mounted located on the test fixture of the vibration shaker. The eigen frequency variation of the thruster module was determined using an accelerometer attached to the main chips located at the center of control board to judge the structural safety of the thruster module before and after full level vibration test. The actual ignition test of the MEMS thruster module was not performed after vibration test because we judged that the thruster module exposed to the vibration loads will be verified through a thermal vacuum test from the design verification point of view. Therefore, the survivability of the electrical function of the thruster module before and after the vibration test was checked by comparing the measured resistance values of the micro-igniter. Figure 11 shows the sine burst test results for the thruster module for each axis. In the sine burst test, a design load of 12 g was applied to the thruster module along each axis by the test method of employing ramp up and down for seven cycles. In addition, sine and random vibration tests were performed according to the qualification test specification described in Table 2. Figure 12 shows the random vibration test results of MEMS thruster module for each axis. The first eigen frequency of the thruster
13
module along each axis was higher than the required value of 120 Hz. The grms values obtained from the random vibration test for each axis were 20.3, 11.7 and 17.8. Table 3 compares the first eigen frequencies obtained from the low level sine sweep (LLSS) test performed before and after each vibration test. The maximum variation of the first eigen frequency of the thruster module was within 1.15% for random vibration test along z-axis, which meets the requirement of the variation being within 5%. Figure 13 shows the measured resistance values of the micro-igniters before and after launch environment tests at the qualification level. The variation of the resistance values were less than 0.5%, and all the functional requirements were successfully verified after the vibration tests. The design and fabrication process of the thruster module, including the modified design on the electrical interface, was successfully qualified, and this can ensure the survivability of the MEMS thruster module in launch environments.
4.2 On-orbit Environment Test Figure 14 shows the thermal vacuum test set-up for the MEMS thruster module to verify the normal operation of the thruster module in a space-simulated thermal vacuum environment. Figure 15 shows the temperature profiles of the thruster module obtained from the thermal vacuum test. The test was performed with low- and hightemperature threshold limits of -35°C and 35°C. The temperature sensors were attached
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to the main chips on the control board and MEMS thruster. The target temperature was measured at the temperature reference point on the MEMS thruster to judge the stabilized temperature. In the test, three repeated cycles of thermal loads were applied to the thruster module, and the actual ignition test of the solid thruster was performed at the lowest temperature condition of -35°C in the last cycle because this is the worst condition from the viewpoint of ignition. To check the state of health of the MEMS thruster during temperature cycling, an electrical function test and a resistance measurement test were performed for the micro-igniters at each hot and cold temperature plateaus before the ignition test at the cold plateau of the last cycle. Thermal balance tests were also conducted at the cold and hot plateaus to construct a highly reliable thermal mathematical model (TMM) of the thruster module by correlating temperature data between test data and simulation results. On-orbit temperature of the MEMS thruster module was predicted again by using the correlated TMM to judge whether the qualification temperature is suitable. Figure 16 indicates the telemetries of the resistance values of the micro-igniters obtained at the hot and cold plateaus of the 1st, 2nd, and 3rd cycles before the thruster ignition test at the last cycle. The variation of the resistance values were less than 0.7% and all functional requirements were also successfully verified in each cycle. Figure 17 shows the condition of the MEMS thruster after the completion of the ignition test. The number on the thruster indicates the order of activation during the test. In the ignition test, five ignitions were activated according to the orbit operation
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scenario. Two micro-igniters positioned along the diagonal direction were activated simultaneously, and one igniter at the center of thruster was activated last to prevent the tumbling of the CubeSat by asymmetric ignition. The test results showed that the six of the nine solid thrusters were successfully activated. The success rate of solid thruster ignition is 66% under vacuum condition. However, a successful ignition of 100% was achieved in the ignition test performed under ambient conditions, though different MEMS thruster samples were used in the test [20]. These observations may be attributed to the differences between vacuum and ambient test conditions rather than to the different test samples because the samples were manufactured by same MEMS fabrication processes. From the test results, we judge that the decrease in success rate under vacuum conditions may be related to the damage to the micro-igniter layer caused by the propagation of micro-cracks after solid thruster ignition under vacuum condition. The vacuum level applied in the test was 10-5 Torr and the condition was more favorable than the ambient test condition for inducing micro-crack propagation. The minimum and maximum prediction temperatures of the MEMS thruster module through on-orbit thermal analysis at satellite level were -29.3°C and 37.5°C, respectively, the worst coldest and hottest orbital conditions. The parameters used in the on-orbit thermal analysis were obtained by correlating the thermal model results with the thermal balance test results; the details, however, are not described in this paper. This means that the qualification test temperatures applied in this study can cover the predicted operating temperature of the mechanism with a margin of 5°C at the lowest
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temperature, though the highest test temperature was slightly greater than the on-orbit predicted temperature. The test results indicated that the MEMS fabrication process of the MEMS solid thruster array was successfully verified through the qualification tests because the thruster module showed a stable function before and after the test. The test results also demonstrated that the design of the electrode pads combined with spring-loaded pins is very suitable for space applications of MEMS technology for ensuring stable electrical contact because the design survived the harsh launch vibration environments. However, the design of the micro-igniter made of Pt electric heating coils fabricated on a 40 mthick glass membrane needs to be modified for ensuring more reliable operation under high-vacuum conditions to avoid damage to the micro-igniter by the propagation of micro-cracks after the ignition of the solid propellant. The flight model of the MEMS solid thruster built by the same fabrication process without modification will be tested again to generate real experimental data and heritage flight achievements on-orbit environments.
5. Conclusion The MEMS-based solid propellant thruster array is one of the main payloads to be verified through the STEP Cube mission. The main purpose of its on-orbit verification is to verify its fabrication process for use in space environment rather than to verify the capability of attitude determination by the MEMS propulsion system. Although the function and performance of the thruster were previously verified by
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laboratory-level research at universities, their design effectiveness has never been qualified in outer space or on-ground simulated space environments. To ensure the successful operation of the MEMS thruster module before flight demonstration on-orbit, launch and on-orbit environment tests were performed at the qualification level. The tests successfully verified the manufacturing process because the thruster module showed stable normal function before the ignition test. The design of electrode pads combined with spring-loaded pins was more reliable than the conventional design. The results indicated that the proposed MEMS fabrication process for space usage proposed in this study has been qualified. However, the design of the micro-igniter shall be modified for ensuring more reliable operation under high vacuum conditions to avoid damage to the micro-igniter by the propagation of micro-cracks after the ignition of the solid propellant. The test results will provide design and fabrication guidelines for MEMS applications for space applications.
Acknowledgments
This research was supported by a research fund (2016) by Chosun University.
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Reference
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Conference on Spacecraft Guidance, Navigation and Control Systems, 2012, p.431. [13] R.A. Zeledon, M.A. Peck, Performance Testing of a CubeSat-scale Electrolysis Propulsion System, in: Proc. of AIAA Guidance, Navigation, and Control Conference, 2012, Paper. AIAA2012–5041. [14] J. Schein, M. Krishnan, J. Ziemer, J. Polk, Adding a "Throttle" to a Clustered Vacuum Arc Thruster, in: Proc. of AIAA NanoTech, At the Edge of Revolution Conference, 2002, Paper. AIAA2002-5716. [15] J. Lee, T. Kim, MEMS solid propellant thruster array with micro membrane igniter, Sensors and Actuators A 190 (2013) 52–60. [16] H.U. Oh, S.H. Jeon, S.C. Kwon, Structural design and analysis of 1U standardized STEP Cube Lab for on-orbit verification of fundamental space technologies, Journal of Materials, Mechanics and Manufacturing 2 (2014) 239–244. [17] H.U. Oh, T.Y. Park, Experimental feasibility study of concentrating photovoltaic power system for cubesat applications, Aerospace and Electronic Systems 51 (2015) 1942–1949. [18] H.U. Oh, M.J. Lee, Development of a non-explosive segmented nut-type holding and release mechanism for cube satellite applications, Transactions of the Japan Society for Aeronautical and Space Sciences 58 (2015) 1–6.
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List of figures Figure 1 Mechanical configuration of the STEP Cube Lab Figure 2 A cross-sectional schematic view of the MEMS thruster (a), the conventional design (b), the modified design (c) Figure 3 Comparison of the cross-sectional temperature contours for direct and indirect heating igniters
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Figure 4 Ignition delay and energy of the direct heating igniter Figure 5 Fabrication process of the micro-igniter Figure 6 The fabricated layer of the MEMS solid propellant thruster and a close-up view of the micro-igniter Figure 7 An exploded view of MEMS thruster module and electrical connections between the thruster and control board Figure 8 Conventional electrical wire soldering method Figure 9 Successful ignition sequence of the MEMS thruster Figure 10 Vibration test set-up Figure 11 Sine burst test result Figure 12 Random vibration test result Figure 13 Changes in the resistance of the micro-igniters before and after the launch vibration tests Figure 14 Thermal vacuum test set-up Figure 15 Thermal vacuum test profile Figure 16 Changes in the resistance of the micro-igniters before and during the thermal vacuum environment Figure 17 Condition of the MEMS thruster after ignition tests in thermal vacuum
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environment
Table 1 Specifications of MEMS-based solid thruster array Description
Specification
Total Mass
5g
Dimensions
30 mm×30 mm×2.5 mm
Propellant
Lead Styphnate (0.02 g)
Burning Time
0.23 ms
Max Thrust
3.62 N
Specific Impulse
62.3 s
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Table 2 Sine & Random Vibration Test Specification Description
Frequency [Hz]
Amplitude
5
1.3 [g]
8
2.5 [g]
100
2.5 [g]
20
0.009 [g2/Hz]
130
0.046 [g2/Hz]
800
0.046 [g2/Hz]
2000
0.015 [g2/Hz]
Overall
8.03 Grms
Sine Vibration
Random Vibration
25
Table 3 LLSS(Low Level Sine Sweep) Test Result Item
Test X Sine Vibration
Y Z
MEMS Thruster Module
X Random Vibration
1st Freq. [Hz]
Axis
Y Z
Before
210.99
After
211
Before
192
After
192
Before
176
After
176
Before
211
After
209
Before
192
After
191
Before
176
After
174
Difference [%] 0 0 0 0.96 0.52 1.15
Highlight
MEMS thruster was selected for the verification test of MEMS technology in space.
The MEMS fabrication methods of thruster were improved for the space
26
application. Launch and on-orbit environment tests were performed at the qualification level.
All the functional requirements were successfully verified after the tests.
27
PII: DOI: Reference:
S0094-5765(16)30629-4 http://dx.doi.org/10.1016/j.actaastro.2016.11.013 AA6077
To appear in: Acta Astronautica Received date: 3 July 2016 Accepted date: 8 November 2016 Cite this article as: Hyun-Ung Oh, Tae-Gyu Kim, Sung-Hyeon Han and Jongkwang Lee, Verification of MEMS Fabrication Process for the Application of MEMS Solid Propellant Thruster Arrays in Space through Launch and Onorbit Environment Tests, Acta Astronautica, http://dx.doi.org/10.1016/j.actaastro.2016.11.013 This is a PDF file of an unedited manuscript that has been accepted for publication. As a service to our customers we are providing this early version of the manuscript. The manuscript will undergo copyediting, typesetting, and review of the resulting galley proof before it is published in its final citable form. Please note that during the production process errors may be discovered which could affect the content, and all legal disclaimers that apply to the journal pertain.
Verification of MEMS Fabrication Process for the Application of MEMS Solid Propellant Thruster Arrays in Space through Launch and On-orbit Environment Tests
Hyun-Ung Oh1, Tae-Gyu Kim1, Sung-Hyeon Han1, and Jongkwang Lee2*
1
Department of Aerospace Engineering, Chosun University, 375 Seosuk-dong, Dong-gu,
Gwangju, 501-759, Republic of Korea 2
Department of Mechanical Engineering, Hanbat National University, 125 Dongseodaero,
Yuseong-gu, Daejeon, 305-719, Republic of Korea
*
Corresponding
author,
[email protected]
ABSTRACT
Tel.:
+82-42-821-1081
Fax:
+82-42-821-1587
e-mail:
One of the most significant barriers encountered to the space application of MEMS technology is its lack of reliability and flight heritage in space environments. In this study a MEMS solid propellant thruster array was selected for the verification test of MEMS technology in space. The function and performance of MEMS solid thruster have been previously verified by laboratory-level research in universities. To ensure the successful operation of the MEMS thruster module before flight demonstration on-orbit, launch and onorbit environment tests were performed at the qualification level. In the launch test, sine burst, and random vibration loads were applied to the MEMS thruster module. The thermal vacuum tests were carried out for the on-orbit environment test. As a result of the launch vibration test and on-orbit environment test, the variations of the characteristics were less than 0.7%, and all the functional requirements were successfully verified after the vibration tests. The tests successfully verified the manufacturing process because the thruster module showed stable normal function before the ignition. The test result outputs will be helpful in establishing MEMS fabrication guidelines for space applications.
Keywords: MEMS Solid thrust, Cube satellite, Launch vibration test, Thermal vacuum test
2
1. Introduction
The mass, volume, and power consumption of an on-board mission payload is a significant contributor to the overall increase in the cost of spacecraft systems. Microelectromechanical systems (MEMS) technology offer several advantages for space applications owing to the drastic reduction in size, weight, and power consumption, made possible by the miniaturization of components and systems [1]. Recently, MEMS is being considered for use in micro-sensors, micro-propulsion systems, and RF switches in space systems for which such devices are of specific interest [2–10]. MEMS systems increase the functionality of satellite systems with limited volume while simultaneously decreasing spacecraft weight. However, to realize the ultimate goal of sending completely micro-fabricated and integrated MEMS-based devices into space, the long-term reliability of MEMS technology needs to be assessed by actual experiments in space, and this will require frequent launches and on-orbit verifications in the near future. Cube satellites (CubeSats) are useful tools for building the space heritage of MEMS technology in on-orbit through orbital experiments and verification tests of new technologies. These satellites are cube-shaped, pico-class miniaturized satellites, and are much smaller than commercial satellites. They usually have a volume of 10 cm3, and a 1
mass of less than 1.33 kg for a standard size of 1U [11]. Recently, CubeSats have been widely advocated by organizations and universities around the world to achieve increasingly
complex
scientific,
surveillance,
interplanetary,
and
technology
demonstration missions. CubeSats capable of orbit control can be used for a wide range of scientific missions; their functionality in an extremely small package promises to yield numerous advanced technologies required for achieving challenging missionrelated functions. In order to realize these missions, new types of micro-thrusters [12– 15] that offer a wide range of impulse bits with much lower total mass are required. The development of micro-thrusters has been actively researched in the recent years, and the performance of various types of thrusters has been investigated. Santoni et al. [12] developed a MEMS-based cold gas micro-propulsion system for attitude control to be tested on board the Ursa Maior CubeSat within the framework of the QB50 projects. The electrolysis propulsion systems [13] investigated by Cornell University are well suited for CubeSat-scale missions because of their simplicity, flexibility in terms of electrical power consumption, and dense propellant storage. These systems provide an estimated 1 km/s of V, which is a significant improvement in the capabilities of CubeSats. A vacuum arc thruster for CubeSat propulsion is also scheduled for an inspace test on board the UIUC CubeSat [14]. Lee et al. [15] proposed an improved
2
MEMS-based solid propellant thruster array with a micro-membrane igniter. They improved the stability of the micro-igniter by using a glass membrane tens of microns thick. They also developed a process to fabricate the micro-igniter on a photosensitive glass wafer instead of on a dielectric membrane. STEP Cube Lab (Cube Laboratory for Space Technology Experimental Project) is the first pico-class satellite being developed at the SSTL (Space Technology Synthesis Laboratory) of Chosun University, and it is scheduled to be launched in 2016 [16]. The main objective of this mission is to perform on-orbit verification of results of spacerelated research conducted at domestic universities. The MEMS-based solid propellant thruster array of Lee et al. [15] is one of the main payloads to be verified through the STEP Cube mission. The main purpose of on-orbit verification of MEMS-based solid propellant thruster arrays is to verify the MEMS fabrication process for use in space environment rather than to verify the capability of the attitude determination by the MEMS propulsion system. This is because the MEMS technology is very attractive for space applications as it offers great benefits such as reduction in size, mass, and power consumption through miniaturization of components and systems. However, to realize the ultimate goal of sending completely micro-fabricated and integrated MEMS systems into space, the long-term reliability of this technology needs to be assessed through
3
experiments in space, and this will require frequent launches in the near future. The MEMS fabrication processes of a solid thruster array such as the sputtering of Pt/Ti heater, diffusion bonding of photosensitive glass, ultraviolet (UV) bonding, solid propellant filling under vacuum condition and electrode design approach will be verified through this program. The successful activation of nine solid thrusters at specified time intervals during life time is used as the success criterion to judge whether the aforementioned MEMS fabrication processes are applicable for space applications. In this paper, we propose a mounting method that employs flexure-like brackets to support the MEMS thruster module, which consists of a MEMS solid thruster and its control board. This method makes it possible to reduce the launch loads transmitted to the thruster module under launch environments. It also allows the thruster module to be easily mounted on or dismounted from the fully integrated satellite structure without disassembling any of the satellite parts during the test and transportation to the launch site. In addition, we propose an electrode pad design combined with spring-loaded pogo-pins instead of soldering of electric wires, which is generally applied in the conventional MEMS fabrication process, to achieve reliable electrical contact between the thruster and its control board under launch and on-orbit environments. In the study, to guarantee successful operation of MEMS thruster on-orbit and to verify the
4
fabrication process including the modified design for space usage, space qualification tests were conducted before flight demonstration. The results obtained from the harsh launch vibration test and thermal vacuum environment test will be helpful for establishing the MEMS fabrication guidelines for space applications.
2. Overview of STEP Cube Lab
Figure 1 shows the mechanical configuration of the STEP Cube Lab; its design is based on a 1U standard CubeSat. The main objectives of the STEP Cube Lab mission [16] are to identify core space technologies studied in domestic industries or universities, and to verify them in orbit. Its primary objective is to perform on-orbit verifications of fundamental space technologies for future space missions. The payloads to be verified in this mission are a variable emittance radiator, a phase change material (PCM), a MEMS-based solid propellant thruster [15], a concentrating photovoltaic (CPV) power system [17], and a novel non-explosive holding and release mechanism triggered by nichrome burn wire heating [18]. The MEMS-based solid propellant thruster located at the bottom of the CubeSat is also one of the main payloads to be verified through the STEP Cube Lab. mission.
5
3. MEMS Solid Propellant Thruster Array
3.1 Design Modification of MEMS Solid Thruster for Space Usage Figure 2 (a) shows a cross-sectional view of the 3×3 MEMS solid propellant thruster array. It consists of four layers of photosensitive glass, a micro-nozzle layer, an intermediary layer, a micro-igniter, a seal layer, and the solid propellant. The micronozzle is located on the first layer, and the second layer is the intermediary layer. The third layer contains the micro-igniter, with a glass membrane tens of microns thick, and the solid propellant. The fourth layer acts as the seal layer. The operation principle is as follows: when the solid propellant attains its ignition temperature achieved by the micro-igniter, the glass membrane is subsequently broken by the high-energy combustion gases of the solid propellant. Thrust is generated through the micro-nozzle. The specifications of the MEMS thruster array are summarized in Table 1. Figure 2 (b) shows the conventional glass-ceramic membrane igniter and Fig. 2 (c) shows the improved glass membrane igniter. In the current design, in order to improve the ignition performance by preventing ignition failures caused by incomplete combustion due to delay in ignition time, the micro-igniter is installed under the glass membrane. Therefore, the solid propellant can be more efficiently ignited because the micro-igniter in the modified design is in direct contact with the solid propellant unlike in the conventional MEMS thruster, where the propellant is heated-up indirectly by an igniter installed on the glass membrane. A numerical simulation of the modified MEMS igniter design for increasing the
6
success rate in orbit was carried out to predict its ignition characteristics using the commercial code, Fluent. The temperature change of the propellant as a function of the time was monitored at 15 V, the required voltage for the ignition of the solid thruster. Figure 3 shows a comparison of the cross-sectional temperature contours between the proposed direct-heating igniter and the conventional indirect-heating igniter at 20ms. The temperature increase rate at the top surface of the propellant was faster for the proposed igniter than for the conventional igniter owing to the position of the heater. The ignition delay and ignition energy of the proposed igniter as a function of the input voltage are presented in Fig. 4. The ignition characteristics depended on the input voltage. The ignition delay and the ignition energy decreased exponentially as the voltage increased. The numerical results showed an ignition delay of 16.8 ms, and the ignition energy of 24.9 mJ at 15 V; these values are about 7.4% lower than the corresponding values for the conventional igniter. In this study, we also proposed a modification of the design of the electrical interface of the MEMS solid thruster to guarantee reliable electrical contact between the thruster and its controller. The electrical interface on the thruster was modified such that it was directly connected to the pogo-pins on the thruster control. The details are described in Section 3.3.
7
3.2 Fabrication Process of the MEMS Solid Propellant Thruster The fabrication process of MEMS thruster was divided into three steps: the fabrication of the micro-nozzle, intermediary, and seal layer; the fabrication of the micro-igniter; and the final assembly process, including the loading of the propellant. The fabrication process of the micro-igniter is shown in Fig. 5. In the first step, the micro-nozzle layer, intermediary layer, and seal layer were fabricated by anisotropic etching of photosensitive glass. The nozzle layer was fabricated by anisotropic etching on one side to obtain the divergence micro-nozzle. The average diameter of nine nozzle throats was 416 μm. The average area ratio of the micro-nozzles was 1.85. The thickness of the micro-nozzle layer, intermediary layer, and seal layer were 0.9 mm, 0.3 mm, and 0.9 mm, respectively. Next, the micro-igniter comprising a heater, a membrane, and a propellant chamber was fabricated. Platinum (Pt) was adopted as the material of the heater because it exhibits high stability at high temperatures and is resistant to oxidation and corrosion. A 40 um thick glass membrane was selected from the viewpoint of structural stability. Glass membranes have lower thermal conductivity and lower fabrication costs than conventional dielectric membranes. The micro-igniter was fabricated by anisotropic wet etching, Pt/Ti deposition and patterning, chemical-mechanical polishing (CMP), and UV bonding process. The glass was then selectively exposed to UV light of wavelength of approximately 310 nm and energy density 2.5 J/cm2. Then, the exposed glass wafer was inserted into a programmable furnace for heat treatment. The glass was heated to
8
500°C at a ramp rate of 2°C /min, and the temperature was maintained constant for 1h. Then, it was heated again to 585°C with a ramp rate of 1°C /min and maintained at this temperature for 1 h. Finally, the glass was cooled at a cooling rat e of −3°C /min. During this heat treatment, the properties of the UV-exposed glass changed because of re-crystallization. After the heat treatment, the wafer was soaked in a diluted 10% hydrofluoric (HF) acid solution. The general etching ratio of glass-ceramic to vitreous glass is 20:1. Because the unexposed area was partially etched, the surface of the wafer was roughened. Subsequently, the wafer was chemically and mechanically polished in order to make its surface flat. The membrane layer was fabricated by patterning the heater and electrodes onto the glass wafer. First, a photoresist was spin-coated onto the glass and patterned by photolithography to form the heater and electrodes. A layer of 200 Å thick Ti layer was sputter deposited to improve Pt adhesion to the glass. A 2000Å thick Pt layer was sputter deposited on the Ti layer. The photoresist on the glass and on the Pt/Ti layer were removed by applying acetone for 2 min. After this process, only the Pt/Ti pattern of the micro-igniter remained. After each layer of the igniter had been fabricated, the membrane layer and the chamber layer were integrated by a UV bonding process. The UV glue was spin-coated onto the membrane layer. The membrane layer and the chamber layer were then aligned, and exposed to UV light of wavelength 310 nm. The fabricated micro-nozzle layer and intermediary layer were polished and cleaned in piranha solution for 10 min. Two layers were pressed together at 500°C by
9
applying a pressure of 2 kPa for 12 h with a ramp rate of 2°C/min and a cooling rate of about −3°C/min. Lead styphnate was selected as the solid propellant, considering its ease of loading and low power consumption for ignition. The solid propellant was dissolved in a solution that contained the solvent and the binder. The solution was put in the chamber layer and dried to remove the solvent. After the filling process, all layers were assembled by UV bonding because the propellant is vulnerable to heat exposure. Figure 6 shows each fabricated layer of the MEMS solid propellant thruster and the fabricated micro-igniter.
3.3 Design of MEMS Solid Thruster Assembly for Space Applications In this study, to achieve a reliable electrical contact between the MEMS thruster and its control board in launch vibration and thermal vacuum environments, an electrode pad design combined with spring-loaded pins instead of the conventional soldering electric wire is proposed. Figure 7 shows an exploded schematic view of the MEMS thruster module and the design of the electrical interface between the thruster and its control board for controlling the activation of the propellants according to the ignition commands on-orbit. The figure also shows the modified electrical interface on the backside of the solid thruster. The conventional design, employing electric wire soldering [15] generally used in the MEMS fabrication process is shown in Fig. 8. However, this approach might not be feasible for space applications considering the harsh launch vibration loads and thermal loads induced by temperature variations on-
10
orbit. Therefore, the conventional design was slightly modified for space applications as follows. To achieve reliable electrical contact and survivability, we proposed electrode pads design combined with spring-loaded pogo-pins. As shown in the exploded schematic view in Fig. 7, the electrical interface on the thruster was directly connected to the pogo-pins on the thruster control board and fixed by a plastic cover made of Delrin® with space heritage. The function of the thruster module was tested by the external command to the thruster control board. The ignition test performed under ambient conditions at the MEMS thruster assembly level showed successful ignition of all nine thrusters. Figure 9 shows the images captured by a high-speed camera, showing the successful ignition of the thruster activated by the control board; the images indicate that the ignition was successful without any malfunction under ambient conditions. In addition, to ensure successful on-orbit verification of the MEMS thruster array, structural design effort at the satellite level are performed to minimize the launch loads transmitted to the thruster under launch environments. This was achieved by mounting the MEMS thruster module on flexure-like brackets made of Delrin® plastic material, as shown in Fig. 1. This design also makes it easier to access and replace the MEMS thruster from outside without de-mating the other parts of the CubeSat because frequent replacements of the thruster are required during the verification test. For example, a dummy thruster without solid propellant will be used to prevent ignition during a full function test of the system level thermal vacuum test. This will also be
11
replaced by a flight model of the thruster with solid propellant before the launch campaign.
4. Space Environmental Qualification Test
One of the most significant barriers encountered to the space usage of MEMS technology is its lack of reliability and flight heritage in space environments. The function and performance of MEMS solid thruster have been previously verified by laboratory-level research in universities, but their design effectiveness has never been qualified in outer space or on-ground simulated space environments. Therefore, it is unclear whether MEMS thruster will operate without any malfunction in an on-orbit environment. In this study, to guarantee the structural safety and successful operation of MEMS thrusters under severe launch and on-orbit environments, we performed a vibration test and thermal vacuum test under qualification level. The main objectives of the qualification tests were to verity the design and fabrication process of the MEMS solid thruster described in the Section 3.2 for judging whether the technology can be used in space applications before actual flight demonstration. The modified design of the electrical interface between the thruster and control board was also verified in the qualification tests. In the vibration test, the effectiveness of the structural design employing a flexure-like bracket design was also investigated.
12
4.1 Launch Vibration Test Figure 10 shows the vibration test set-up for the qualification test of the MEMS thruster module. In the test, sine burst, and random vibration loads were applied to the MEMS thruster module. The qualification test level was based on the QB50 test specifications summarized in Table 2 [19]. An accelerometer sensor, which is used to apply input test loads was mounted located on the test fixture of the vibration shaker. The eigen frequency variation of the thruster module was determined using an accelerometer attached to the main chips located at the center of control board to judge the structural safety of the thruster module before and after full level vibration test. The actual ignition test of the MEMS thruster module was not performed after vibration test because we judged that the thruster module exposed to the vibration loads will be verified through a thermal vacuum test from the design verification point of view. Therefore, the survivability of the electrical function of the thruster module before and after the vibration test was checked by comparing the measured resistance values of the micro-igniter. Figure 11 shows the sine burst test results for the thruster module for each axis. In the sine burst test, a design load of 12 g was applied to the thruster module along each axis by the test method of employing ramp up and down for seven cycles. In addition, sine and random vibration tests were performed according to the qualification test specification described in Table 2. Figure 12 shows the random vibration test results of MEMS thruster module for each axis. The first eigen frequency of the thruster
13
module along each axis was higher than the required value of 120 Hz. The grms values obtained from the random vibration test for each axis were 20.3, 11.7 and 17.8. Table 3 compares the first eigen frequencies obtained from the low level sine sweep (LLSS) test performed before and after each vibration test. The maximum variation of the first eigen frequency of the thruster module was within 1.15% for random vibration test along z-axis, which meets the requirement of the variation being within 5%. Figure 13 shows the measured resistance values of the micro-igniters before and after launch environment tests at the qualification level. The variation of the resistance values were less than 0.5%, and all the functional requirements were successfully verified after the vibration tests. The design and fabrication process of the thruster module, including the modified design on the electrical interface, was successfully qualified, and this can ensure the survivability of the MEMS thruster module in launch environments.
4.2 On-orbit Environment Test Figure 14 shows the thermal vacuum test set-up for the MEMS thruster module to verify the normal operation of the thruster module in a space-simulated thermal vacuum environment. Figure 15 shows the temperature profiles of the thruster module obtained from the thermal vacuum test. The test was performed with low- and hightemperature threshold limits of -35°C and 35°C. The temperature sensors were attached
14
to the main chips on the control board and MEMS thruster. The target temperature was measured at the temperature reference point on the MEMS thruster to judge the stabilized temperature. In the test, three repeated cycles of thermal loads were applied to the thruster module, and the actual ignition test of the solid thruster was performed at the lowest temperature condition of -35°C in the last cycle because this is the worst condition from the viewpoint of ignition. To check the state of health of the MEMS thruster during temperature cycling, an electrical function test and a resistance measurement test were performed for the micro-igniters at each hot and cold temperature plateaus before the ignition test at the cold plateau of the last cycle. Thermal balance tests were also conducted at the cold and hot plateaus to construct a highly reliable thermal mathematical model (TMM) of the thruster module by correlating temperature data between test data and simulation results. On-orbit temperature of the MEMS thruster module was predicted again by using the correlated TMM to judge whether the qualification temperature is suitable. Figure 16 indicates the telemetries of the resistance values of the micro-igniters obtained at the hot and cold plateaus of the 1st, 2nd, and 3rd cycles before the thruster ignition test at the last cycle. The variation of the resistance values were less than 0.7% and all functional requirements were also successfully verified in each cycle. Figure 17 shows the condition of the MEMS thruster after the completion of the ignition test. The number on the thruster indicates the order of activation during the test. In the ignition test, five ignitions were activated according to the orbit operation
15
scenario. Two micro-igniters positioned along the diagonal direction were activated simultaneously, and one igniter at the center of thruster was activated last to prevent the tumbling of the CubeSat by asymmetric ignition. The test results showed that the six of the nine solid thrusters were successfully activated. The success rate of solid thruster ignition is 66% under vacuum condition. However, a successful ignition of 100% was achieved in the ignition test performed under ambient conditions, though different MEMS thruster samples were used in the test [20]. These observations may be attributed to the differences between vacuum and ambient test conditions rather than to the different test samples because the samples were manufactured by same MEMS fabrication processes. From the test results, we judge that the decrease in success rate under vacuum conditions may be related to the damage to the micro-igniter layer caused by the propagation of micro-cracks after solid thruster ignition under vacuum condition. The vacuum level applied in the test was 10-5 Torr and the condition was more favorable than the ambient test condition for inducing micro-crack propagation. The minimum and maximum prediction temperatures of the MEMS thruster module through on-orbit thermal analysis at satellite level were -29.3°C and 37.5°C, respectively, the worst coldest and hottest orbital conditions. The parameters used in the on-orbit thermal analysis were obtained by correlating the thermal model results with the thermal balance test results; the details, however, are not described in this paper. This means that the qualification test temperatures applied in this study can cover the predicted operating temperature of the mechanism with a margin of 5°C at the lowest
16
temperature, though the highest test temperature was slightly greater than the on-orbit predicted temperature. The test results indicated that the MEMS fabrication process of the MEMS solid thruster array was successfully verified through the qualification tests because the thruster module showed a stable function before and after the test. The test results also demonstrated that the design of the electrode pads combined with spring-loaded pins is very suitable for space applications of MEMS technology for ensuring stable electrical contact because the design survived the harsh launch vibration environments. However, the design of the micro-igniter made of Pt electric heating coils fabricated on a 40 mthick glass membrane needs to be modified for ensuring more reliable operation under high-vacuum conditions to avoid damage to the micro-igniter by the propagation of micro-cracks after the ignition of the solid propellant. The flight model of the MEMS solid thruster built by the same fabrication process without modification will be tested again to generate real experimental data and heritage flight achievements on-orbit environments.
5. Conclusion The MEMS-based solid propellant thruster array is one of the main payloads to be verified through the STEP Cube mission. The main purpose of its on-orbit verification is to verify its fabrication process for use in space environment rather than to verify the capability of attitude determination by the MEMS propulsion system. Although the function and performance of the thruster were previously verified by
17
laboratory-level research at universities, their design effectiveness has never been qualified in outer space or on-ground simulated space environments. To ensure the successful operation of the MEMS thruster module before flight demonstration on-orbit, launch and on-orbit environment tests were performed at the qualification level. The tests successfully verified the manufacturing process because the thruster module showed stable normal function before the ignition test. The design of electrode pads combined with spring-loaded pins was more reliable than the conventional design. The results indicated that the proposed MEMS fabrication process for space usage proposed in this study has been qualified. However, the design of the micro-igniter shall be modified for ensuring more reliable operation under high vacuum conditions to avoid damage to the micro-igniter by the propagation of micro-cracks after the ignition of the solid propellant. The test results will provide design and fabrication guidelines for MEMS applications for space applications.
Acknowledgments
This research was supported by a research fund (2016) by Chosun University.
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Reference
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Conference on Spacecraft Guidance, Navigation and Control Systems, 2012, p.431. [13] R.A. Zeledon, M.A. Peck, Performance Testing of a CubeSat-scale Electrolysis Propulsion System, in: Proc. of AIAA Guidance, Navigation, and Control Conference, 2012, Paper. AIAA2012–5041. [14] J. Schein, M. Krishnan, J. Ziemer, J. Polk, Adding a "Throttle" to a Clustered Vacuum Arc Thruster, in: Proc. of AIAA NanoTech, At the Edge of Revolution Conference, 2002, Paper. AIAA2002-5716. [15] J. Lee, T. Kim, MEMS solid propellant thruster array with micro membrane igniter, Sensors and Actuators A 190 (2013) 52–60. [16] H.U. Oh, S.H. Jeon, S.C. Kwon, Structural design and analysis of 1U standardized STEP Cube Lab for on-orbit verification of fundamental space technologies, Journal of Materials, Mechanics and Manufacturing 2 (2014) 239–244. [17] H.U. Oh, T.Y. Park, Experimental feasibility study of concentrating photovoltaic power system for cubesat applications, Aerospace and Electronic Systems 51 (2015) 1942–1949. [18] H.U. Oh, M.J. Lee, Development of a non-explosive segmented nut-type holding and release mechanism for cube satellite applications, Transactions of the Japan Society for Aeronautical and Space Sciences 58 (2015) 1–6.
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[19] F. Singarayar, R. Reinhard, C. Asma, J. Thoemel, T. Scholz, C. Bernal, W. Weggelaar, G. Shirville, D. Kataria, M. Richard, QB50 system requirements and recommendations issue 4 (2013) www.qb50.eu. [20] D. Seo, J. Lee, S. Kwon, The development of the micro-solid propellant thruster array with improved repeatability, Journal of Micromechanics and microengineering 22 (2012) 094004.
List of figures Figure 1 Mechanical configuration of the STEP Cube Lab Figure 2 A cross-sectional schematic view of the MEMS thruster (a), the conventional design (b), the modified design (c) Figure 3 Comparison of the cross-sectional temperature contours for direct and indirect heating igniters
22
Figure 4 Ignition delay and energy of the direct heating igniter Figure 5 Fabrication process of the micro-igniter Figure 6 The fabricated layer of the MEMS solid propellant thruster and a close-up view of the micro-igniter Figure 7 An exploded view of MEMS thruster module and electrical connections between the thruster and control board Figure 8 Conventional electrical wire soldering method Figure 9 Successful ignition sequence of the MEMS thruster Figure 10 Vibration test set-up Figure 11 Sine burst test result Figure 12 Random vibration test result Figure 13 Changes in the resistance of the micro-igniters before and after the launch vibration tests Figure 14 Thermal vacuum test set-up Figure 15 Thermal vacuum test profile Figure 16 Changes in the resistance of the micro-igniters before and during the thermal vacuum environment Figure 17 Condition of the MEMS thruster after ignition tests in thermal vacuum
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environment
Table 1 Specifications of MEMS-based solid thruster array Description
Specification
Total Mass
5g
Dimensions
30 mm×30 mm×2.5 mm
Propellant
Lead Styphnate (0.02 g)
Burning Time
0.23 ms
Max Thrust
3.62 N
Specific Impulse
62.3 s
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Table 2 Sine & Random Vibration Test Specification Description
Frequency [Hz]
Amplitude
5
1.3 [g]
8
2.5 [g]
100
2.5 [g]
20
0.009 [g2/Hz]
130
0.046 [g2/Hz]
800
0.046 [g2/Hz]
2000
0.015 [g2/Hz]
Overall
8.03 Grms
Sine Vibration
Random Vibration
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Table 3 LLSS(Low Level Sine Sweep) Test Result Item
Test X Sine Vibration
Y Z
MEMS Thruster Module
X Random Vibration
1st Freq. [Hz]
Axis
Y Z
Before
210.99
After
211
Before
192
After
192
Before
176
After
176
Before
211
After
209
Before
192
After
191
Before
176
After
174
Difference [%] 0 0 0 0.96 0.52 1.15
Highlight
MEMS thruster was selected for the verification test of MEMS technology in space.
The MEMS fabrication methods of thruster were improved for the space
26
application. Launch and on-orbit environment tests were performed at the qualification level.
All the functional requirements were successfully verified after the tests.
27