Launch Loads Research Articles

Design of an adjustable bipod flexure for a large-aperture mirror of a space camera.

An adjustable bipod flexure (ABF) technique for a large-aperture mirror of a space camera is presented. The proposed flexure mount can decrease the surface distortions caused by the machining error and the assembly error of the mirror assembly (MA) in a horizontal optical testing layout. Through the analysis of the compliance matrix of conventional bipod flexure, the positional relationship between the rotation center and the apex of the flexure is investigated. Then, the principle of the adjustable flexure, known as the trapezoidal switching principle, is proposed based on the analysis result. The structure and application of the flexure are also described. The optical performance of the mirror mounted by the adjustable flexures in different misalignments was performed using finite element methods. The result shows that the astigmatic aberration due to gravity is effectively reduced by adjusting the mount, and the root-mean-square value of the mirror can be minimized with the misalignment between the flexure pivot and the neutral plane minimized. New monolithic bipod flexures, based on the optimal regulating variable Δu according to the measurement results, are manufactured to replace the ABFs to secure the mirror's safety against launch loads. Modal analysis verified the mechanical safety of the MA with respect to the new monolithic flexures.

Universal CubeSat Platform Design Technique

This article summarizes cubesat technology, provides examples of their scientific impact, and describes the design and the manufacturing of a Cubesat platform. As for the design of the overall frame structure of the CubeSat, we have searched a lot of literature and consulted many predecessors' designs, and collected many satellite structure images. After analyzing the data, we aimed at all kinds of different structures’ advantages and disadvantages, finally we got a best design. It is a satellite of cubic shape (10 cm per side), weighing approximately 1kg, based on the creation of a central body made of different material using the 3D-Printing techniques. The 3D-Printing technique has several advantages including fast implementation, accuracy in manufacturing small parts and low cost. Moreover, concerning the construction of a small satellite, this technique is very useful thanks to the accuracy achievable in details, which are sometimes difficult and expensive to realize with the use of tools machine. The structure must be able to withstand the launch loads. For this reason, several simulations using an FEM simulation and an intensive vibration test campaign will be performed in the system development and test phase.

Utilization of trash for radiation protection during manned space missions

High-energy charged particles in space pose a severe threat to the health and safety of astronauts. This is especially the case in deep space exploration owing to the absence of magnetic field protection as present in low Earth orbit (LEO) manned spaceflight. This has necessitated the investigation and development of effective space radiation protection materials and methods. Manned space missions produce a significant amount of trash, which can be compressed by a heat melt compactor (HMC) to reduce space utilization. The trash would also be sterilized during the compaction process and the extracted water can be recycled. The processed trash can potentially be used for space radiation protection, with the benefit of reduced launch load. In this study, the Monte Carlo method was used to acquire information about the primary and secondary radiation particles that emerged from different radiation shields made from an HMC-processed model trash of a manned space mission, as well as those made from aluminum and water, which are common, and currently used space radiation protection materials. The types and energies of the considered incident radiation particles were based on space radiation environment spectra. A comparison of the space radiation protection capacities of the different shields revealed that HMC-processed trash was superior to aluminum and water. Processed trash thus promises to be a practicable alternative to water for the construction of radiation emergency areas for future deep space missions as the mission proceeds and the water is consumed.

Combined Instrumentation Package COMARS+ for the ExoMars Schiaparelli Lander

In order to measure aerothermal parameters on the back cover of the ExoMars Schiaparelli lander the instrumentation package COMARS+ was developed by DLR. Consisting of three combined aerothermal sensors, one broadband radiometer sensor and an electronic box the payload provides important data for future missions. The aerothermal sensors called COMARS combine four discrete sensors measuring static pressure, total heat flux, temperature and radiative heat flux at two specific spectral bands. The infrared radiation in a broadband spectral range is measured by the separate broadband radiometer sensor. The electronic box of the payload is used for amplification, conditioning and multiplexing of the sensor signals. The design of the payload was mainly carried out using numerical tools including structural analyses, to simulate the main mechanical loads which occur during launch and stage separation, and thermal analyses to simulate the temperature environment during cruise phase and Mars entry. To validate the design an extensive qualification test campaign was conducted on a set of qualification models. The tests included vibration and shock tests to simulate launch loads and stage separation shocks. Thermal tests under vacuum condition were performed to simulate the thermal environment of the capsule during the different flight phases. Furthermore electromagnetic compatibility tests were conducted to check that the payload is compatible with the electromagnetic environment of the capsule and does not emit electromagnetic energy that could cause electromagnetic interference in other devices. For the sensor heads located on the ExoMars back cover radiation tests were carried out to verify their radiation hardness. Finally the bioburden reduction process was demonstrated on the qualification hardware to show the compliance with the planetary protection requirements. To test the actual heat flux, pressure and infrared radiation measurement under representative conditions, aerothermal tests were performed in an arc-heated wind tunnel facility. After all qualification tests were passed successfully, the acceptance test campaign for the flight hardware at acceptance level included the same tests than the qualification campaign except shock, radiation hardness and aerothermal tests. After passing all acceptance tests, the COMARS+ flight hardware was integrated into the Schiaparelli capsule in January 2015 at the ExoMars integration site at Thales Alenia Space in Turin. Although the landing of Schiaparelli failed, resulting in the loss of most COMARS+ flight data because they were stored on the lander, some data points were directly transmitted to the orbiter at low sampling rate during the entry phase. These data indicate that all COMARS+ sensors delivered useful data until parachute deployment with the exception of the plasma black-out phase. Since measured structure and sensor housing temperatures are far below predicted pre-flight values, a new calibration using COMARS+ spare sensors at temperatures below 0 °C is necessary.

On the development of Antenna feed array for space applications by additive manufacturing technique

Space agencies are looking for advanced technologies to build light weight and stiff payload components to bear space environment and launch loads. Additive manufacturing (AM) techniques like Direct Metal Laser Sintering (DMLS) is one of the suitable option which can be explored for space applications. This paper highlights the development process of Antenna Feed Array (AFA) using DMLS as an Additive Manufacturing (AM) technique. A high efficiency horn element is used in the array. Such horns are preferred for this development as they are the prime choice for feed elements in High Throughput satellites (HTS) that employ Multibeam Antennas. A brief description of Multibeam antennas along with the RF design process for the high efficiency horn is presented. In the development process, certain design rules of AM are adopted based on consideration to produce self-sustaining structures. AFA realized by DMLS is evaluated by functional testing, vibration testing for space qualification test levels. Test results shows its structural intactness which proves its space worthiness. Procedures are very well established for further development of space payload components.

Thermal Load Minimization of the Frequency Domain Multiplexed Readout for the Athenar X-IFU Instrument

The high sensitivity of cryogenic TES-based detectors opens new windows for astrophysical observations ranging from (far) infrared to X-rays. A number of operational and future space and ground-based instruments rely on cryogenic detectors to improve their performance with respect to the capabilities of earlier technologies. To reach the required sensitivities, base temperatures as low as 50 mK are necessary, and stringent requirements on magnetic shielding, microvibrations, and temperature stability are applicable. To minimize the heat load and complexity of the instruments, efficient multiplexing schemes and low-power amplifiers are needed. In addition, for space-based cryogenic instruments, mechanical launch loads and power consumption limitations constrain the available parameter space for engineering further. This paper discusses the system design considerations which are applicable to optimize the multiplex factor within the boundary conditions as set by the space craft for the X-IFU instrument on the Athena observatory. More specifically, the interplay between the science requirements such as pixel dynamic range, pixel speed, and cross talk, and the space craft requirements such as the power dissipation budget, and available bandwidth will be discussed.

Slipway Pile Footing Stiffness and Transient Launch Loads

This paper presents information on the necessity of accounting for the effect of transient ship launch loads on inclined longitudinal slipway structures when evaluating the design stiffness of the pile footing. A considerable increase in the stiffness is registered in comparison with such structure's resistance under static loads conditions. The introduction of an additional factor of operating condition, to increase the design bearing capacity of piles under instantaneous loads is substantiated.

Prospect for UV observations from the Moon. II. Instrumental design of an ultraviolet imager LUCI

We present a design for a near-ultraviolet (NUV) imaging instrument which may be flown on a range of available platforms, including high-altitude balloons, nanosatellites, or space missions. Although all current UV space missions adopt a Ritchey-Chretain telescope design, this requires aspheric optics, making the optical system complex, expensive and challenging for manufacturing and alignment. An all-spherical configuration is a cost-effective and simple solution. We have aimed for a small payload which may be launched by different platforms and we have designed a compact, light-weight payload which will withstand all launch loads. No other UV payloads have been previously reported with an all-spherical optical design for imaging in the NUV domain and a weight below 2 kg. Our main science goal is focussed on bright UV sources not accessible by the more sensitive large space UV missions. Here we discuss various aspects of design and development of the complete instrument, the structural and finite-element analysis of the system performed to ensure that the payload withstands launch-load stresses and vibrations. We expect to fly this telescope -- Lunar Ultraviolet Cosmic Imager (LUCI) -- on a spacecraft to the Moon as part of the Indian entry into Google X-Prize competition. Observations from the Moon provide a unique opportunity to observe the sky from a stable platform far above the Earth's atmosphere. However, we will explore other opportunities as well, and will fly this telescope on a high-altitude balloon later this year.

Thermo-mechanical Design for On-orbit Verification of MEMS based Solid Propellant Thruster Array through STEP Cube Lab Mission

A MEMS solid propellant thruster array shall be operated within an allowable range of operating temperatures to avoid ignition failure by incomplete combustion due to a time delay in ignition. The structural safety of the MEMS thruster array under severe on-orbit thermal conditions can also be guaranteed by a suitable thermal control. In this study, we propose a thermal control strategy to perform on-orbit verification of a MEMS thruster module, which is expected to be the primary payload of the STEP Cube Lab mission. The strategy involves, the use of micro-igniters as heaters and temperature sensors for active thermal control because an additional heater cannot be implemented in the current design. In addition, we made efforts to reduce the launch loads transmitted to the MEMS thruster module at the system level structural design. The effectiveness of the proposed thermo-mechanical design strategy has been demonstrated by numerical analysis.

큐브위성 탑재를 위한 MEMS 고체 추력기의 구조설계 및 검증

MEMS solid thruster module is composed of solid thruster and its control board. It was developed for the purpose of an academic research. Therefore, thermo-mechanical design and verification for space usage were not considered in the design phase. To mount it on a cube satellite without any design modification, technical efforts at the system level structure design is required. In this study, we proposed a structural design concept to mount the MEMS thruster module by using brackets for guaranteeing structure safety under launch loads and easier mating and de-mating of MEMS thruster module during test phase. The effectiveness of the design has been verified through structural analysis and vibration test. In addition, electrical connection method using spring pins between MEMS thruster and control board is effective for guaranteeing the structural safety under launch vibration loads.

Conducting Wall Hall Thrusters

A unique configuration of the magnetic field and channel geometry near the wall of Hall thrusters, called magnetic shielding, has recently demonstrated the ability to significantly reduce the erosion of the boron nitride (BN) walls and extend the life of Hall thrusters by orders of magnitude. The ability of magnetic shielding to minimize interactions between the plasma and the discharge chamber walls in regions where erosion typically occurs has for the first time enabled the replacement of insulating walls with conducting materials without loss in Hall thruster performance. It is important to note that this is not a thruster with anode layer (TAL) where the walls are at or near cathode potential, but is a Hall thruster configuration where the walls are near the anode potential. The BN rings in the 6-kW H6 Hall thruster were replaced with graphite that self-biased to near the anode potential during operation. The thruster efficiency remained over 60% (within 2% of the baseline BN configuration) with a small decrease in thrust and increase in Isp typical of magnetically shielded Hall thrusters. The graphite wall temperatures decreased significantly compared with both shielded and unshielded BN configurations, leading to the potential for higher power operation. Eliminating ceramic walls makes it simpler and less expensive to fabricate a thruster to survive launch loads, and the graphite discharge chamber radiates more efficiently, which increases the power capability of the thruster compared with conventional Hall thruster designs.

Flight Performance of a Small Diameter Munition with a Rotating Wing Actuator

Future enhanced lethal effects at the infantry squad level likely include precision guided technologies. The focus of this study is maneuvering projectiles launched from man-portable weapon systems. A novel guided projectile concept is proposed for achieving control authority requirements in the challenging environment of low-dynamic pressure, small size, high launch loads, spin stabilization, and low cost. This new maneuver concept is based on a rotating wing actuator. Experimental and advanced computational aerodynamics techniques were applied. Aerodynamic models and projectile flight mechanics were derived to enable flight simulation. Assessment of ballistic delivery accuracy, based on physics-based models of the delivery process, was undertaken to quantify control authority requirements. Maneuvering flight simulations demonstrated that this concept affords enough course correction to compensate for ballistic delivery errors.

Hypervelocity Impact Phenomenon in Bulk Metallic Glasses and Composites**

Collisions with debris are major cause of concern for spacecraft and satellites. Developing new materials that can combat these threats, while still providing low‐density and sufficient toughness to survive launch loads, is important for future spacecraft design. In the current work, hypervelocity impacts are used to estimate the ballistic limit for bulk metallic glass and their composites and to investigate spalling behavior. The composites are shown to have excellent combinations of hardness and toughness for use as shields.

Dynamic characterization of satellite assembly for responsive space applications

The rapid deployment of satellites for responsive space surveillance applications is hindered by the need to flight-qualify their components and the resulting mechanical assembly. Conventional methods for qualification testing of satellite components are costly and time consuming. Furthermore, full-scale vehicles must be subjected to simulated launch loads during testing, and this harsh testing environment increases the risk of damage to satellite components during qualification. This work focuses on replacing this potentially destructive testing procedure with a non-destructive structural health monitoring (SHM)-based technique while maintaining the same level of confidence in the testing procedure's ability to qualify the satellite for flight. We focus on assessing the performance of SHM techniques to replace the high-cost qualification procedure and to localize faults introduced by improper assembly. The goal of this work is to create a dual-use system that can both assist in the process of qualifying the satellite for launch, as well as provide continuous structural integrity monitoring during manufacture, transport, launch and deployment. SHM techniques were applied on a small-scale structure representative of a responsive satellite. The test structure consisted of an extruded aluminum space-frame covered with aluminum shear plates assembled using bolted joints. Multiple piezoelectric transducers were bonded to the test structure and acted as combined actuators and sensors. Piezoelectric active-sensing based techniques, including measurements of low-frequency global frequency response functions and high-frequency wave propagation techniques, were employed. Using these methods in conjunction with finite element modeling, the dynamic properties of the test structure were established and areas of potential damage could be identified and localized. A procedure for guiding the effective placement of the sensors and actuators is also outlined.

Optomechanical analysis of a 1-m light-weight mirror system

We present the optomechanical analysis results for a 1-m light-weight mirror system for a space telescope. The mirror has partially-closed pockets at the back surface and three square bosses at the rim for flexure mounting. The mirror design is optimized to satisfy the performance requirements under launch loads and a space environment. Mirror surface distortions due to gravity, assembly, isothermal loads, and moisture absorption are investigated. Mechanical safety issues such as mirror fracture, flexure yielding, and adhesive breakage are also examined.

A passive launch and on-orbit vibration isolation system for the spaceborne cryocooler

To enhance the pointing performance by isolating disturbances induced by a spaceborne payload mounted cryogenic cooler during the on-orbit operation, a passive launch and on-orbit vibration isolation system has been proposed and investigated. The isolation system is also effective for reducing the transmitted launch loads to the cooler. In order to demonstrate the effectiveness of the isolation system, launch environment tests such as sine, random and shock tests and a micro-vibration measurement test of the cooler combined with the isolator have been performed. The test results indicate that the isolation system guarantees both the structural safety of the cooler under the launch environment and the micro-vibration isolation of the cooler under the on-orbit condition.

Adjustable bipod flexures for mounting mirrors in a space telescope

A new mirror mounting technique applicable to the primary mirror in a space telescope is presented. This mounting technique replaces conventional bipod flexures with flexures having mechanical shims so that adjustments can be made to counter the effects of gravitational distortion of the mirror surface while being tested in the horizontal position. Astigmatic aberration due to the gravitational changes is effectively reduced by adjusting the shim thickness, and the relation between the astigmatism and the shim thickness is investigated. We tested the mirror interferometrically at the center of curvature using a null lens. Then we repeated the test after rotating the mirror about its optical axis by 180° in the horizontal setup, and searched for the minimum system error. With the proposed flexure mount, the gravitational stress at the adhesive coupling between the mirror and the mount is reduced by half that of a conventional bipod flexure for better mechanical safety under launch loads. Analytical results using finite element methods are compared with experimental results from the optical interferometer. Vibration tests verified the mechanical safety and optical stability, and qualified their use in space applications.

Plastic Cubesat: An innovative and low-cost way to perform applied space research and hands-on education

This paper describes the design and the manufacturing of a Cubesat platform based on a plastic structure.The Cubesat structure has been realized in plastic material (ABS) using a “rapid prototyping” technique. The “rapid prototyping” technique has several advantages including fast implementation, accuracy in manufacturing small parts and low cost. Moreover, concerning the construction of a small satellite, this technique is very useful thanks to the accuracy achievable in details, which are sometimes difficult and expensive to realize with the use of tools machine. The structure must be able to withstand the launch loads. For this reason, several simulations using an FEM simulation and an intensive vibration test campaign have been performed in the system development and test phase. To demonstrate that this structure is suitable for hosting a complete satellite system, offering innovative integrated solutions, other subsystems have been developed and assembled.Despite its small size, this single unit (1U) Cubesat has a system for active attitude control, a redundant telecommunication system, a payload camera and a photovoltaic system based on high efficiency solar cells.The developed communication subsystem has small dimensions, low power consumption and low cost. An example of the innovations introduced is the antenna system, which has been manufactured inside the ABS structure. The communication protocol which has been implemented, the AX.25 protocol, is mainly used by radio amateurs. The communication system has the capability to transmit both telemetry and data from the payload, in this case a microcamera.The attitude control subsystem is based on an active magnetic system with magnetorquers for detumbling and momentum dumping and three reaction wheels for fine control. It has a total dimension of about 50×50×50mm. A microcontroller implements the detumbling control law autonomously taking data from integrated magnetometers and executes pointing maneuvers on the basis of commands received in real time from ground.The subsystems developed for this Cubesat have also been designed to be scaled up for larger satellites such as 2U or 3U Cubesats. The additional volume can be used for more complex payloads. Thus the satellite can be used as a low cost platform for companies, institutions or universities to test components in space.

PRO/MECHANICA-based structural and random vibration analysis of picosatellite structure

This paper focuses on the structural and vibration analysis of a low-cost picosatellite structure to the launch load environment. This picosatellite complies with the CubeSat program design specifications, where the satellite is less than one kilogram in mass with a dimension of 10 cm3. The satellite structure is treated as a combination of beams and thin plate elements. A detailed finite element model is developed to observe its responses to the quasi-static accelerations induced by the launch vehicle. The natural frequencies of the prototype were computed numerically through finite element based methods and compared with the natural frequencies induced by the launch vehicle. It was also found that the maximum stress experienced by the structure during launch was well within the stress limits of the material.

Vibroacoustic launch analysis and alleviation of lightweight, active mirrors

Lightweight, active, silicon carbide mirrors can increase the capability of space-based optical systems. However, launch survival is a serious concern for such systems, with the vibrations and acoustics from launch threatening to damage the optics. Therefore, a dynamic, state-space launch model has been developed with which one can quickly analyze the survival probability of many designs and also directly analyze launch load alleviation techniques. This paper discusses the launch model from which launch stress and survival probability are obtained, as well as launch load alleviation techniques that may increase the probability of launch survival. Three launch load alleviation techniques are presented and analyzed: whole spacecraft isolation, passive shunt circuits using the existing embedded actuators, and active damping using the existing actuators. All of the techniques reduce the launch stress, but at the expense of mass and complexity. The launch model allows for early identification of lightweight, active mirror designs which will survive launch, and the alleviation techniques expand the feasible design space by decreasing the launch stress and increasing the probability of launch survival.