Ultimate load and release time controllable non-explosive separation device using a shape memory alloy actuator

Transcription

1 Journal of Mechanical Science and Technology 25 (5) (2011) 1141~ DOI /s Ultimate load and release time controllable non-explosive separation device using a shape memory alloy actuator Wonjun Tak, Minsu Lee and Byungkyu Kim * Korea Aerospace University, 200-1, Hwajeon-dong, Deokyang-gu, Goyang-si, Gyeonggi-do, Korea (Manuscript Received October 6, 2010; Revised January 5, 2011; Accepted January 19, 2011) Abstract This paper describes a newly designed ultimate load and release time controllable non-explosive separation device which is activated by a spring-type shape memory alloy (SMA) actuator. This device is comprised of a separation mechanism consisting of a deformation module, a blocker, the housing, two release springs and a spring-type SMA actuator. Through a theoretical approach, the ultimate load of the separation device and the required force to deform the deformation module according to the thickness of the deformation module are investigated. Based on theoretical results, the specifications of an SMA actuator to deform the deformation module are determined. In order to validate theoretical result, we manufactured a separation device which has a 1mm thickness of deformation module. Subsequently, the release time test, preload test, ultimate load test, and shock test are performed respectively. As a result, the release time with 30 W of power input is 55 sec. It actuates reliably under 150 N of preload and generates a maximum shock of G when the diameter of the SMA wire is 1.75mm. The maximum ultimate-load for 1mm thickness of aluminum deformation module is 1,510N. Conclusively, satellite designers can select proper specification such as the ultimate load, shock level and release time according to satellite requirements by simply changing the thickness of the deformation module. Keywords: Non-explosive actuator; Separation device; SMA actuator; Solar panel; Deployment Introduction This paper was recommended for publication in revised form by Associate Editor Jeong Sam Han * Corresponding author. Tel.: , Fax.: address: bkim@kau.ac.kr KSME & Springer 2011 In order for a satellite mission to start successfully, the satellite must be separated on time from the launcher in space. After the satellite gets in the target orbit, appendices such as solar panels and antennas should be separated for power generation and communication. In the early stages of the development of the satellite, the pyrotechnic-based separation devices are widely used because of a simple operating principle and fast response time. However, it requires extreme care in handling, since the pyrotechnic-based separation device is comprised of gunpowder [1]. In addition, it generates a pyroshock over 5,000 g, as well as releasing contaminants, since the device is activated by the ignition of the gunpowder. Therefore, it sometimes causes malfunctions in electrical devices and dirt particles on the optical lenses or solar panels in the satellite. There have been many failed satellite missions because of the pyro-technique separation device. A survey of pyroshock flight failures revealed 83 shock related anomalies out of 600 launches with over 50% of these resulting in catastrophic failure [2]. Recently, the rapid growth of nano-technology has enabled small satellites to have almost all the functions that big and medium-sized satellites have. From a cost effectiveness viewpoint, these small satellites have many advantages compared to the big and medium-sized satellites. As a satellite's size decreases, many devices within the satellite should be tightly integrated. As a result, satellites have become more sensitive to shocks and contaminants. Therefore, the pyrotechnic-based separation devices, which generate high shock and contaminants, are not suitable for small satellites [2]. There has been a great deal of research activity for developing a non-explosive separation device [3, 4]. In another attempt to replace the pyrotechnic-based separation device, we designed and manufactured a non-pyrotechnic separation device that is activated by a spring-type SMA actuator. The proposed device comprises a deformation module, blockers, the housing, two release springs, and a spring-type SMA actuator. Since the deformation module is deformed by the activation of a spring-type SMA actuator, we could control the ultimate load and release time by simply changing the thickness of the deformation module. In the first step of design, the thickness of the deformation module, which can satisfy various satellite ultimate loads, is determined by finite element analysis (FEA) and the

2 1142 W. Tak et al. / Journal of Mechanical Science and Technology 25 (5) (2011) 1141~1147 Table 1. Material of each component. Part Block Housing Aluminum- deformation- module Blocker Spring type SMA actuator Release spring Material Aluminum (Al) NiTi Shape memory alloy Steel spring Fig. 2. Proposed separation device. Fig. 1. Configuration of separation device. result is confirmed experimentally. Then, a spring-type SMA actuator is designed, and the response time of the SMA actuator is studied under normal power of a small satellite. In order to obtain the space heritage, the ultimate load test, shock test, and response time test, all of which are important factors to evaluate the separation device, are performed respectively. 2. Device design and operating 2.1 Device design The proposed separation device is designed to hold a satellite appendage (e.g. solar panel) with high binding force and generate a low shock in the release sequence. As shown in Figs. 1 and 2, it consists of the deformation module, the housing, two release springs, a block, an SMA actuator and blockers. All materials for each component (Table 1) that are space proven are utilized [5]. 2.2 Working principle The main purpose of the separation device is to bind the appendage to the main body of the satellite before the mission and release it on time by command. In order to generate a low shock for the satellite, the separation device needs to be actuated by a small force. Therefore, the spring-type SMA actuator is employed in this device. Fig. 3 shows the working principle of the proposed separation device. At the initial stage (1), the deformation module, which combines with a block, is blocked by the blocker that is welded on the housing. In order to keep contact between the deformation module and the blocker, two release springs are installed under the block. Fig. 3. Working principle. An SMA actuator is installed in the deformation module under elongated conditions. On the operating stage, as the electric current flows through the SMA, the SMA actuator is heated. Consequently, the spring-type SMA actuator retracts and causes plastic deformation of the deformation module (2). Then, the deformation module is released from the blocker. Finally, the block is released from the housing (3) when the release springs are elongated, and it is pushed out from the housing. 2.3 Design of the non-explosive separation device Due to the diversity of the required loads depending on the specific satellite appendage, there are various requirements for the ultimate load and release time. Therefore, we invented a non-explosive separation device that could satisfy the requirements of the ultimate load and release time according to the mission of each satellite. In the proposed device, the ultimate load and release time can be controlled by adjusting the thickness of the deformation module. After deciding on the appropriate thickness of the deforma-

3 W. Tak et al. / Journal of Mechanical Science and Technology 25 (5) (2011) 1141~ Fig. 4. Relationship between ultimate load, thickness of deformation module and required force to deform the deformation module. tion module for the ultimate load requirement of a satellite, the dimensions of the SMA actuator are designed. Then, the response time of the SMA actuator is calculated by controlling the input power. Through this sequence, the proposed device can satisfy the various release times and ultimate loads Design of the deformation module Based on the finite element analysis (FEA) using the commercial code of ANSYS, the force which can deform the deformation modules made of aluminum is investigated. Accordingly, the ultimate load is obtained using FEA and the result is confirmed by an experimental result using a Material Test System (MTS-810 from MTS Korea ) as shown in Fig. 4. In order to confirm the selection of the ultimate load and release time by controlling the thickness of the deformation module, a prototype with a thickness of 1 mm that can be deformed by 58 N of SMA actuator force is fabricated. In that case, 1,510 N of the ultimate load can be guaranteed as shown in Fig Design of SMA actuator As previously mentioned, 58 N of force is required to deform the deformation module with a thickness of 1 mm. Therefore, a spring-type SMA actuator that can generate the force to deform is designed in consideration of the space limitations. Based on a calculation of the design parameters, a springtype actuator with an outer diameter of 10 mm is manufactured using an SMA wire-diameter of 1.75 mm. The design procedure is as follows: First, using Eq. (1), spring rate (K) is determined [6]. 4 3 Gd K = (1) 8nD Considering the spring length and safety margin, the number of turns and spring stroke are decided. Second, through the basic equation between the deflection and the spring constant, Fig. 5. Generative force according to displacement. the relationship between force and deflection can be determined. p δh = Kh δ = δ + S l h From the limitation of module dimension, S is calculated. Then, a 19 mm spring-type SMA actuator is manufactured, which is fully compressed and in the shape memorized condition. When the SMA actuator is installed on the deformation module, it is extended to 28 mm. As power is engaged, the spring-type SMA actuator that has a wire-diameter of 1.75 mm generates a sufficient force of 60 N to deform the aluminum deformation module with a thickness of 1 mm when it returns to the shape memorized condition. The relationship between displacement and generative force obtained using Eqs. (1) and (2) is presented in Fig. 5. Through the experiments under configuration of Fig. 6, the calculated value of the generative force according to displacement is validated as in Fig. 5. Also, the calculated generative force of 2 mm and 1.5 mm wire diameters of SMA is also presented for a non-explosive separation device that requires other specifications such as ultimate load and response time. The actuation time of the SMA actuator is one of the most important design parameters for reliable non-explosive separation devices. Therefore, the Joule heating method was selected to activate the SMA actuator, which was reliable. We performed the response time test for the separation device comprised with spring-type SMA actuator with an SMA wire-diameter of 1.75 mm under 10-3 torr of vacuum condition and 7.2 A x 4.2 V of input power. It was measured as 55 seconds. Then, the theoretical analysis on the SMA actuation time is carried out with Eq. (3). When we calculate the SMA actuation time using the Runge-kutta method, it is 52.5 seconds. The small difference between the calculated value and experimental value is regarded as the minor convection effect in the vacuum chamber and mechanism operating time. (2)

4 1144 W. Tak et al. / Journal of Mechanical Science and Technology 25 (5) (2011) 1141~1147 Table 2. Design procedure from ultimate load to actuation time of proposed separation device which has SMA actuator as 19mm of fully compressed length and 10mm of outer-diameter. Design Procedure Fig. 6. Experiment configuration to measure generative force Ultimate load (N) Deformable aluminium SMA wire diameter (mm) module thickness (mm) Exerting force of SMA actuator (N) Actuation time (sec) , , , time from 13.8 s to 87.2 s and have an ultimate load from 676 N to 2,524 N as shown in Fig. 7. Therefore, the proposed device can be developed to satisfy a diversity of satellite specifications. 3. Performance test For space applications, the reliability of the device is the most important factor. In order to confirm reliability, functional tests are carried out. Among the test requirements for space heritage, the ultimate load test, response time test, and shock-level test are performed. Fig. 7. Variation of response time and ultimate load according to SMA wire diameter e π εσ sur 2 dt IR' d ( T T ) = (3) dt ρc( πd /4) Based on the above bench marking, which is a comparison of the experimental and calculation results, we can estimate the SMA response time. Accordingly, ultimate load can be estimated for various diameters without experiments. Table 2 shows the design procedure in detail. Once the requirement for ultimate load (1) is determined, the thickness of the deformation module (2) can be calculated using FEA analysis. Then, we can design the dimensions (diameter, spring length, stroke, etc.) of the spring-type SMA actuator (3) that can generate the force to deform the deformation module (4). Accordingly, the actuation time (5) can be determined using Eq. (3). In summary, we can achieve a faster response time with the thin deformation module but lose the ultimate load that guarantees safety of the device structure. If a high ultimate load is required for any satellite, we should consider the proposed device if the slow response time is allowable. Conclusively, with an assembly of thickness ( mm) of the deformation module and diameter (1-2 mm) of SMA wire in the proposed device, it can activate with a response 3.1 Determination of the axial load In the developing stage of the satellite, the maximum tensile force, which the separation device should withstand, is determined based on a quasi-static load according to the launching environment. The quasi-static load is the acceleration load that is exerted on the mass center of the satellite. Therefore, every part has a different quasi-static load according to the shape of the satellite. In other words, there is no specific value for quasi-static loads because the value can be changed depending on the size of the satellite, the number of folding solar panels and many other variables. Based on the launching experience equation, the quasi-static load from a relationship between mass (m) and acceleration (a) can be calculated [7]. 0.3 = 50/ [ ] (4) a m G From Eq. (4), if we assume a solar panel has 5 kg of mass, it will have 30.8 G of quasi-static acceleration and consequently, it will be under 1,509 N of quasi-static load. Therefore, the ultimate load test for the developed device is performed at 1,509 N. As mentioned previously, some parameters such as the mass of the solar panel, safety factors, and binding points are assumed for this specific case. Therefore, we need to determine the quasi-static load according to the design of each satellite. Accordingly, the load requirement can be satisfied through the control of the deformation module of the proposed separation device.

5 W. Tak et al. / Journal of Mechanical Science and Technology 25 (5) (2011) 1141~ tests and shock tests are carried out under the defined preload. Shock tests are performed on aluminum panels with a thickness of 6 mm, which are assumed to be used in the satellite body Preload generator To measure the shock level, the preload that comes from the weight of solar panels should be engaged to simulate the actual satellite situation. Adding dummy mass is a common way to generate the preload. However, the preload generator should be able to reduce momentum from gravity and be designed to change the preload easily. Therefore, the preload generator is designed using action and reaction phenomena. Fig. 8 shows the principle for the preload generator [8]. Fig. 8. Design and concept of preload generator. Fig. 9. Shock test setup configuration System configuration for shock level test The separation device will be in a zero G state when it operates in the space environment. The test environment is simulated by using bungee cables to suspend the device in order to emulate the rigid body motion in space. Also, the satellite body is made of 500x500x6 mm aluminum panel. The preload of 150N, which is the maximum guaranteed preload (proof load) of the proposed separation device, is set by using the preload generator. The accelerometers are installed at 30 mm (Pos. 1), 60 mm (Pos. 2) and 90 mm (Pos. 3) respectively, from the installed position of the separation device. The shock test set-up and measurement set-up on the panel are shown in Figs. 9 and Shock level test result In general, the explosive separation device generates a shock level of over 5,000 G and the shock level of a nonexplosive separation device is required to be under 500 G. In comparing those requirements, the very low shock from the proposed non-explosive separation device is generated. Fig. 11 in the time domain and Fig. 12 in the frequency domain show the results of the shock level test under 150 N of preload. It generates a maximum of G at Pos. 1 (30 mm) from the installed point of the separation device. Without the preload, it only generates a maximum of G on the Pos. 1 (30 mm) [9]. Compared with other developed separation devices, the proposed separation device shows very low levels of acceleration and frequency [10]. 4. Future work Fig. 10. Shock measurement setup on panel. 3.2 Shock test Generating a low shock level is the biggest advantage for this non-explosive separation device. To perform the shock test, the preload that comes from between the satellite body and solar panel should be defined first. Then, performance The proposed separation device will not operate in the ordinary atmosphere, but in a space environment which has a high vacuum level and extreme temperatures. Also, when the launcher goes up to the orbit, the separation device has to endure extreme vibrations and secure the appendage safely. Therefore, we plan to perform thermal-vacuum testing referring to European Space Agency (ESA) standardization [11]. The proposed device will be tested in the environment of the operating temperature range from -60 to 70 at a maximum vacuum level of 10-5 torr as pre-developed SMA separation devices have done [3]. Also, through the vibration

6 1146 W. Tak et al. / Journal of Mechanical Science and Technology 25 (5) (2011) 1141~ Conclusion Fig. 11. Shock level test result in time domain. We developed an ultimate load and response time controllable non-explosive separation device which is activated by a spring-type shape memory alloy (SMA) actuator. It is achieved by assembling a deformation module, blockers, release springs and a spring-type SMA actuator. Since the deformation module is deformed by the activation of a spring-type SMA actuator, we could control the ultimate load and response time by simply changing the thickness of the deformation module. As the thickness of the deformation module increases, a higher ultimate load could be obtained. However, the response time increased since we spent more time to heat up the springtype SMA actuator with a larger wire diameter. With one prototype model, we could confirm that the proposed device could stand 1,510 N of ultimate load. The shock level, an important factor for a small satellite, was measured as G under 150 N of preload. With 30 W of operation power, the response time was 55 seconds. Based on the working principle of the proposed device, we expect to manufacture a non-explosive low shock separation device that is customized to the specifications of ultimate load and response time. Acknowledgment This work has been supported by the National Research Foundation of Korea (NRF) through the National Space Laboratory Project ( ). Fig. 12. SRS plot result in frequency domain. Fig. 13. Concept of the separation device with redundant function. test and other required tests such as creep, external shock, life cycle [3], we will verify the proposed separation device. In addition, combining two proposed separation devices serially as shown in Fig. 13, we will add redundant function to the proposed device and perform the additional tests to obtain space heritage. Nomenclature d : SMA wire diameter G h : High temperature modulus G l : Low temperature modulus n : Number of turns D : Spring diameter δ h : High temperature deflection δ l : Low temperature deflection p : Generative force of shape memory alloy S : Defined stroke I : Input current R e : Resistance of SMA ε : Emissivity σ : Stefan-Boltzmann number T : Aimed temperature (=85 ) T sur : Surround temperature (=25 ) ρ : Density of SMA c : Specific heat References [1] J. H. Lim, K. W. Kim, S. W. Kim, C. H. Lee, J. H. Rhee and D. S. Hwang, Non-Explosive Actuator Technology for Satellite Applications, KARI (2009)

7 W. Tak et al. / Journal of Mechanical Science and Technology 25 (5) (2011) 1141~ [2] E. R. Fosness, S. J. Buckley, Waylon and F. Gammill, Deployment and release devices efforts at the air force research laboratory space vehicles, AIAA Space (2001) Conference and Exposition, AIAA [3] [4] [5] T. P. Sarafin, Spacecraft structures and mechanism from concept to launch, Space Technology Library, USA (2007). [6] T. Waram, Actuator design using shape memory alloys, Ontario L8S 1L8, Canada, second edition (1993) [7] T. S. Jang, H. B. Kim, S. H. Woo, S. H. Lee and M. R. Nam, Analysis on environmental test specifications for dollar panels of STSAT-2, The Korean Society for Noise and Vibration Engineering (KSNVE) (2005) [8] J. Műller and C. Zauner, Low shock release unit-easy resettable and 100% reusable, Proceedings of the 10th European Space Mechanisms and Tribology Symposium, 24-26, September [9] I. Kipp, PSD AND SRS IN SIMPLE TERMS, ISTA Con 98, Orlando, FL (1998). [10] NAS , Aerospace Systems Pyrotechnic Shock Data (Ground Test and Flight), Vol. VI. Goddard Space Flight Center : NASA, USA (1970). [11] ECSS-E-10-03A, Space engineering (Testing), ECSS Secretariat. Requirements & Standard Division, Noordwijk, Netherlands (2002). Minsu Lee received a B.A. degree in mechanical engineering from Korea Aerospace University in His research interests include micro actuators using smart materials and micro control systems. Byungkyu Kim received his PhD. in mechanical engineering from the University of Wisconsin, Madison, in From 1997 to 2000, he was a technical staff member of CXrL (Center for X-ray Lithography) in the University of Wisconsin where he developed a computer code for thermal modeling of a mask membrane and wafer during beam exposure. From 2000 to 2005, he worked for Microsystem Center of KIST as a principal research scientist. He was in charge of developing the microrobot for a microcapsule-type endoscope. Currently, he is an associate professor in the school of aerospace and mechanical engineering in Korea Aerospace University. His research interests include microelectromechanical systems (MEMS) actuators and smart actuator based space mechanism. Wonjun Tak graduated from Korea Aerospace University and is now a graduate student in Korea Aerospace University. His major interest is mechanical designs, space applications and space qualifications.