Launch Vibration Environment Research Articles

Design of a New Multi-Variant Solar Panel Deployment System (Mvspds)

Multi-Variant Solar Panel Deployment System (MVSPDS) is defined as the satellite deployment System which can change its orientation according to the power supply required for the satellite or power supply source available to the satellite. To deploy the solar panels completely, it is necessary to design the deployment mechanism which has high precision and reliability. So, this method of deployment will not only provide the said benefit but also it will allow a wide range of application. Consequently, the analysis on the dynamic characteristic of the deployment mechanism must be done at an initial design stage. The design effectiveness and structural safety of the proposed solar panel module were validated by launch vibration and in-orbit environment tests at the qualification level. In this paper, the complete design of a new Multi-Variant Solar Panel Deployment System in a Satellite is proposed completely by me, where the author have inculcated various deployment methods and proposed a new method of satellite deployment. The complete design is done in the Autodesk 360 software (Educational license) where the design and animation method are extensively used to make this possible. It will be clearly depicted from the design that the structure is very compact during the process of orbit insertion. Yet when it come into working it definitely works according to the power need of the satellite. For an example when the satellite needs less amount of power it will transform itself in such a way that the amount of power generation will be less on the other hand when the satellite needs more power than the Solar Panel will change itself in such a way that it will have the maximum surface area as a result of which the power generation will be maximum. Moreover, the variety which provides is really unique as it also has the power to change its solar panel into different structure for an example it can change itself into a circle like structure it can change itself in the canister to form another structure and moreover it will work completely according to the need if programmed properly for space flight mission. And hence can give rise to a wide variety of different deployment methods.

Thermo-Mechanical Design and Validation of Spaceborne High-Speed Digital Receiver Unit for Synthetic Aperture Radar Application

This paper presents the effectiveness of the thermo-mechanical design of a high-speed digital receiver unit (HSDRU) developed for spaceborne synthetic aperture radar applications. The main features of HSDRU’s thermo-mechanical design include the thermal management of high-heat dissipation units by adopting heat sinks with the additional function of structural stiffeners and securing the heat rejection path to the upper side of electronics that interfaces the foil radiator for the on-orbit passive thermal control of electronics exposed to deep space environments. The thermal design, which adopts a thermal gap pad, is closely related to the solder joint fatigue life in a launch vibration environment, owing to its initial compressive static load between the heat sink and heat dissipation units that enhances the heat transfer capability. The effectiveness of the design was validated via the qualification level of launch environment tests.

Development of a Novel Deployable Solar Panel and Mechanism for 6U CubeSat of STEP Cube Lab-II

The structural safety of solar cells mounted on deployable solar panels in the launch vibration environment is a significant aspect of a successful CubeSat mission. This paper presents a novel highly damped deployable solar panel module that is effective in ensuring structural protection of solar cells under the launch environment by rapidly suppressing the vibrations transmitting through the solar panel by constrained layer damping achieved using printed circuit board (PCB)-based multilayered thin stiffeners with double-sided viscoelastic tapes. A high-damping solar panel demonstration model with a three-pogo pin-based burn wire release mechanism was fabricated and tested for application in the 6U CubeSat “STEP Cube Lab-II” developed by Chosun University, South Korea. The reliable release function and radiation hardness assurance of the mechanism in an in-orbit environment were confirmed by performing solar panel deployment tests and radiation tests, respectively. The design effectiveness and structural safety of the proposed solar panel module were validated by launch vibration and in-orbit environment tests at the qualification level.

CubeSat’s Deployable Solar Panel with Viscoelastic Multilayered Stiffener for Launch Vibration Attenuation

Ensuring the structural safety of a deployable solar panel under a severe launch vibration environment is one of the important factors for a successful CubeSat mission. A CubeSat’s deployable solar panel proposed in this study is effective to guarantee the structural safety of solar cells by attenuating launch loads owing to the superior damping characteristic achieved by a multilayered stiffener with viscoelastic acrylic tapes. The demonstration model of 3 U CubeSat’s deployable solar panel was fabricated and tested to validate the effectiveness of the proposed design. The basic dynamic characteristics of the solar panel were measured through free-vibration tests according to the various layers of the stiffener. Moreover, the characteristics of the deployed solar panel were measured and investigated under various temperatures to predict its capability under in-orbit operation. The effectiveness of the proposed design for launch vibration attenuation was demonstrated through qualification level sine and random vibration tests.

Research on Anti-Vibration based on Screw Fixation and Friction-Assisted Fixation Method for the Space Payloads

Rack is kind of the experimental platform that used in the space station for performing scientific experiment. Many payloads are mounted on Rack with limited operating space. The shipment of a Rack to its final destination in orbit is far more complicated than for all other commercial products. The launch vehicle-induced vibration and shock environment can be a payload killer. Based on the limited operating space, this paper proposes a screw fixation and friction-assisted fixation method. Through establishing the force and the parameters of the screw fixation and friction-assisted fixation based on the payload, deriving the mathematical model and carrying out the dynamic environment test. Test results match theoretical results quite well. The screw fixation and friction-assisted fixation method can effectively resist excessive dynamic and shock loads.

Development of 6 U CubeSat’s Deployable Solar Panel with Burn Wire Triggering Holding and Release Mechanism

In the present work, a deployable solar panel based on a burn wire triggering holding and release mechanism was developed for use of 6 U CubeSat. The holding and release mechanism was designed based on a nichrome burn wire cutting method widely used for CubeSat applications. However, it provides a high loading capability, reliable wire cutting, multiplane constraints, and handling simplicity during the tightening process of wire. A demonstration model of a printed circuit board-based solar panel stiffened by a high-pressure fiberglass-laminated G10 material was fabricated and tested to validate the effectiveness of the design and functionality of the mechanism under various test conditions. The structural safety of the solar panel combined with the mechanism in a launch vibration environment was verified through sine and random vibration tests at qualification level.

적층형 블레이드가 적용된 냉각기용 진동절연기의 발사환경에서의 구조건전성 검증

우주용 냉각기는 관측위성의 적외선 검출기 초점면부의 극저온 냉각을 위해 적용되며, 궤도운용 시 단일 주파수 형태의 미소진동을 발생한다. 상기 미소진동은 관측위성의 영상품질을 저하시키는 주된 요인이며, 이에 따라 미소진동 절연이 요구된다. 본 연구에서는 상기 우주용 냉각기의 미소진동 절연을 목적으로 별도의 발사구속장치 적용 없이도 발사진동 및 궤도 미소진동환경에 동시적용 가능한 적층형 블레이드 진동절연기를 제안하였다. 본 진동절연기에 적용된 블레이드는 얇은 금속 박판을 다층으로 적층하고 각 층 상호면에 점탄성 특성 부여가 가능한 테이프를 적용하여 발사환경에서의 피로 내구성 향상 및 고댐핑 특성 부여를 목적하였다. 제안된 진동절연기의 기본특성 확인을 위해 자유감쇠 시험을 실시하였으며, 인증수준에서의 발사진동 시험을 통해 설계유효성을 입증하였다.

Evaluation of structural design methodologies for predicting mechanical reliability of solder joint of BGA and TSSOP under launch random vibration excitation

The fatigue damage that accumulates on a solder joint under a launch vibration environment is a major cause of failure in spaceborne electronics. Thus, predicting the mechanical safety of the solder joint in the structural design phase is important to guarantee a successful space mission. This paper proposes a mechanical design methodology to predict and ensure more reliable mechanical safety of the solder joints of a plastic ball grid array (PBGA) and a thin-shrink small outline package (TSSOP) under launch random vibration excitation. Random vibration fatigue tests for a sample printed circuit board (PCB) assembly were performed to assess the fatigue life of BGA and TSSOP solder joints. The effectiveness of the proposed structural design methodologies was evaluated by comparing the margin of safety calculated from various analysis approaches and the random vibration fatigue test results of the PCB specimen.

히트싱크 적용에 따른 BGA 패키지의 발사진동 및 궤도 열환경 조건에서의 피로수명의 해석적 검토

히트싱크 적용에 따른 BGA 패키지의 발사진동 및 궤도 열환경 조건에서의 피로수명의 해석적 검토

Micro-jitter attenuation of spaceborne cooler by using a blade-type hyperelastic shape memory alloy passive isolator

In this study, the primary design objective is to develop a passive isolator that can guarantee structural safety of the cooler assembly in a launch vibration environment without a launch locking mechanism, while effectively isolating the cooler-induced micro-jitter during the on-orbit operation of the cooler. To achieve the design objective, we focused on the utilization of characteristics of the hyperelastic shape memory effects. The major advantage of the isolator is that the micro-jitter isolation performance is much less sensitive to the aligned position of the isolator in comparison with the conventional isolator. Moreover, implementation of an additional 0g compensation device during a satellite level on-ground test, such asa jitter measurement test, is not required. In this study, the basic characteristics of the isolator were measured using the torque test and free vibration test. The micro-jitter attenuation capability and position sensitivity of the proposed isolator design were validated by the micro-jitter measurement test.

Experimental Design Validation of Tilting Calibration Mechanism by Using Shape Memory Alloy Spring Actuator

A tilting calibration mechanism is periodically deployed to view the reference temperature target during on-board calibration of a spaceborne imaging sensor and stowed after calibration. In the present work, we have proposed a new design strategy using a shape memory alloy (SMA) spring as an actuator that provides a fail-safe function to prevent the blocking of the main optical path when the mechanical driving part of the mechanism is stopped at a certain position during on-board calibration. Although a launch locking device was not considered in the design, this approach makes it possible to impose mechanical constraints on the driving part of the mechanism in severe launch vibration environments. The effectiveness of the proposed design was experimentally validated by a deploying and stowing function test and launch vibration environment tests such as a sine burst test, a random vibration test, and a pyroshock simulating impulse shock test. The test results demonstrated that the mechanism fulfills all the required functions for on-board calibration. The use of an SMA spring actuator was proved effective for implementing the dual function of a fail-safe in an emergency phase and a mechanical constraint on the driving part of the mechanism in severe launch vibration environment.

Experimental study on on-orbit and launch environment vibration isolation performance of a vibration isolator using bellows and viscous fluid

Operation of various actuators on a satellite such as reaction wheels, cryocoolers, etc. produces micro-vibration which may have adverse effects on the performance of high precision payloads. In order to protect the high precision payloads from the vibration disturbances, vibration isolators are often applied in the vibration propagation path. Vibration isolators developed for on-orbit space application should be able to isolate micro-vibration when in orbit, and also withstand severe launch vibration environment. In this paper, on-orbit and launch environment isolation performance of a vibration isolator using bellows and viscous fluid is evaluated. For the on-orbit environment test, a test setup was prepared inside a thermal-vacuum chamber and the transmissibility of the vibration isolator was measured to examine the effects of on-orbit space environment on the isolation performance of the developed isolator. The isolation test performed in a vacuum condition revealed problems caused by the elongation of the bellows due to the pressure difference which were addressed by sealing the viscous fluid inside the bellows at a vacuum condition. The effect of temperature variation on the isolation performance was found to be confined to the narrow frequency region around the isolator's resonant frequency due to the isolator's three parameter configuration. This result implies that the isolation performance of a vibration isolator having temperature dependent damping can be made less sensitive to the temperature variation by implementing three parameter configuration. The developed vibration isolator is designed to survive large deformation without being permanently deformed to guarantee the structural safety of the isolator in the launch environment. The structural safety of the isolation system is evaluated through sine and random vibration and shock tests. The on-orbit and launch environment test results indicate that the developed isolator can provide effective isolation for the small amplitude vibration disturbances on-orbit and also survive the large amplitude vibration during launch without any damage.

Performance evaluation of spaceborne cryocooler micro-vibration isolation system employing pseudoelastic SMA mesh washer

A spaceborne cryocooler produces undesirable micro-vibration disturbances during its on-orbit operation, which is one of the main sources of degradation of the image quality of high-resolution observation satellites. Therefore, to comply with the strict mission requirement for the acquisition of high-quality images, micro-vibration disturbances induced by cryocooler operation need to be isolated. In this study, we proposed a spaceborne cryocooler micro-vibration isolator that employs a pseudoelastic shape memory alloy (SMA) mesh washer, which guarantees vibration isolation performance in a severe launch vibration environment while effectively isolating the micro-vibrations from the cryocooler on-orbit. Basic characteristics of the cryocooler assembly integrated with the proposed isolators were measured through static tests and free vibration tests. The effectiveness of the isolator design was demonstrated by the micro-vibration measurement tests under qualification temperature limits.

A study of joint damping in metal plates

For satellite applications the determination of the correct dynamic behaviour and in particular the structural damping is important to assess the vibration environment for the spacecraft subsystems and ultimately their capability to withstand the launch vibration environment. Therefore, the object of this investigation is to experimentally analyse a range of aluminium panel configurations to study the effect of joints on the damping of the complete structure. The paper begins with a full description of the experimental method used to accurately determine the modal loss factors for each of the panel configurations analysed. Nine different panels were used in the experimental tests, six of which incorporate lap joints variations. The joint parameters investigated include fastener type, bolt torque, fastener spacing, overlap distance and the effect of stiffeners. The damping results of ten different joint variants are presented for each of the first twelve modes of vibration. This data is directly compared to the damping factors of an equivalent monolithic panel. Various specific conclusions are made with respect to each of the joint parameters investigated. However, the primary conclusion is that the mode shape combined with the joint stiffness and joint location can be suggestive as to the likely magnitude increase of the modal loss factor.

Television Camera System for the Command Module of the Apollo Spacecraft

The Apollo mission to the moon will be covered by television. This paper describes the lightweight, handheld TV camera that will be used by the Apollo astronauts. The camera's mission requirements, parameters, and system description are discussed. The special considerations in the selection of major designs, such as sensor, optics, power supply, and automatic light-level compensation are given. In addition, a detailed theory of operation of the video amplifier, yoke and deflection circuits, beam current regulator, and sync generator is presented. Since a prime requisite for the camera is light weight, hybrid and monolithic integrated circuits are used extensively in the camera electronics. A description of these circuits is given. Factors such as dynamic range of illumination, operation during a launch vibration environment, and handheld-sighting requirements with a narrow-angle zoom lens are discussed. The trade-offs between frame rate, resolution, and bandwidth relevant to image smear and image quality are also considered for this particular application.