Orbiter Payload Research Articles

Design of Orbital Payload transfer Vehicle

Design of Orbital Payload transfer Vehicle

Prediction Aerospace Software to Detect Solar Activity and the Fast Tracking of Solar Activity Using Remote Sensing Instruments in Orbit

At present, solar remote sensing instruments face problems regarding the short detection time for solar activity and the need to preheat the electronics before detection. The accurate orbit prediction function ensures that the maximum error of the payload electronics system preheating time is less than 8 s, enabling the electronics to obtain stable and reliable solar remote sensing data, and allowing the prediction function to adapt to the changes in preheating time and the limit position of turntable rotation. To achieve the complete detection of solar activity by a payload in orbit, a turntable with two directions of rotation was accurately controlled to achieve rapid capture and track the Sun’s activity, and the delay time for tracking the Sun successfully was less than 8 s to ensure that the payload made full use of the solar detection time for full detection. The turntable’s tracking error was less than 0.002°, allowing the payload to obtain high-precision solar remote sensing data in orbit. The above methods enable the payload to realize the complete, stable, and high-precision detection of solar activity in orbit for wavelengths ranging approximately from 165 nm to 2400 nm.

In Situ Propellant Alternatives for a Lunar Ascent/Descent Vehicle

Recently, NASA has pushed for returning humans to the moon sustainably with in situ resource utilization as the central focus. The moon has an abundance of water that is proposed to be electrolyzed into hydrogen and oxygen to be used as propellant. Other volatiles such as ammonia, carbon dioxide, and methane are also present. A mission architecture for a lunar ascent/descent vehicle (LADV) from the Polytechnic University of Turin and nuclear thermal propulsion (NTP) engine models from the University of Alabama in Huntsville were used to compare in-situ-derived propellants for a LADV. This study considered a LADV originating from the lunar surface, delivering a payload in the lunar distant retrograde orbit, and returning to the lunar surface for retanking. This research analyzed the impacts on this mission of using hydrogen NTP, water/ammonia NTP, liquid-oxygen augmented nuclear thermal rocket, and Aeon 1 methane–oxygen engines using the selected architecture and tools. The results were compared to the reference hydrogen–oxygen RL10 engine. The propulsion system comparison analysis showed that combustion engines will offer better overall performance than NTP-based engines due to a 50% decrease in propellant volume, a 20% decrease in dry mass, and a lower propellant mass than the water and ammonia NTP systems. Both the hydrogen–oxygen and methane–oxygen propulsion systems will have similar propellant masses when compared to other systems. This is due to the order of magnitude higher mass of the NTP engines, with the highest mass contribution coming from the reactor. However, both water and ammonia alternative propellant NTP engines can still be viable candidates for the usage of these minimally processed propellants to satisfy this mission.

Determining the degree of effect of heat flows on the deformation of the shell of a space inflatable platform with a payload

This paper reports a study into the influence exerted by the thermal flows of space environment on the deformation of the shell of a space inflatable platform with a payload. The mathematical model of the effect of temperature fluctuations on the mass-inertial characteristics of the space inflatable platform of an ellipsoidal shape has been improved. The following assumptions were introduced to the model. The temperature distribution on the illuminated part and the unlit part of the shell is uniform. The gradient of the temperature difference between the illuminated and unlit parts is the same for all points of the shell. To determine deformations, a moment-free theory was used. The model of the space inflatable platform is a «rubber bullet» that works only for stretching and compression. All deformations are elastic. The advantages and limitations of the use of the developed mathematical model have been determined. Computer simulation of the orbital motion of a space inflatable platform with a payload in a sun-synchronous orbit was carried out. The material of the platform shell is Kapton. Estimates of temperature fluctuations in the illuminated and unlit part of the shell and the temperature of the gas inside it were obtained. The dependence of elastic deformations on temperature was determined, taking into account the Young’s modulus of the material. The influence of changes in gas pressure on the movement of payload attachment points and the change in the inertia tensor have been determined. The obtained results showed that the inertia tensor varies within the order of 10–5 kgm2. The maximum deviation of the fastening points of the payload from the initial position on the illuminated part of the shell was about 10–6 m. Considering the stability of the structure to the effects of heat flows of the space environment, the possibility of using space inflatable platforms as a means for separating a grouping of satellites has been shown

MOVE-III: A CubeSat for the detection of sub-millimetre space debris and meteoroids in Low Earth Orbit

The Munich Orbital Verification Experiment (MOVE) is a CubeSat student project housed at the Scientific Workgroup for Rocketry and Spaceflight at the Technical University of Munich. MOVE-III is the fourth CubeSat under development, and the first 6U mission of the MOVE project that will carry a dedicated scientific payload in orbit. The mission aims at acquiring in-situ observations of sub-millimetre space debris and meteoroids in Low Earth Orbit, with the objective of compiling a dataset of flux, as well as object mass and velocity measurements that can be used for the validation of the small object estimates of space debris models and support further studies related to the characterisation of the space environment. The MOVE-III CubeSat employs the MOVE-BEYOND platform and is planned to carry three Debris Density Retrieval and Analysis (DEDRA) plasma ionization sensors. The Preliminary Design Review has been completed in early 2022, with the next milestone being the Critical Design Review, planned for 2023. The paper elaborates on the scientific objectives of the mission and the expected data products, provides an overview of the detector operation principle and presents the overall system architecture, the platform configuration and the subsystem interaction. Considerations on the debris mitigation aspects of the mission are additionally discussed.

CubeSat project: experience gained and design methodology adopted for a low-cost Electrical Power System

This paper focuses on experimented space projects run by universities, offering an effective design process to improve learning methods in space engineering. The approach used for the design of CubeSat Electrical Power System (EPS) will go through estimations, sizing, simulations, PCB design and end with an experimental test procedure for design validation. The main design criteria presented in this paper are low costs and effective reliability. To meet the first criterion, the Commercial-Off-The-Shelf (COTS) components are used during the design, which has become an effective path to put experimental payloads in orbit for minimal cost. In the second criterion, the function of some critical EPS components was replicated, such as circuits used for power maximization and battery charge regulation, by following a suitable algorithm embedded in MCU and hot/cold redundant analog integrated circuits. However, the choice was made on the state of the art of CubeSat components already tested. This approach of working will help new space engineers to think about optimal solutions for the design of the appropriate EPS for a CubeSat university project achievable on a limited budget. Finally, the necessary experiments of the designed EPS were carried out and the results were illustrated in the context below.

Orbit-Injection Strategy and Trajectory-Planning Method of the Launch Vehicle under Power Failure Conditions

Aiming at the problem of autonomous decision making and trajectory planning (ADMTP) for launch vehicles under power failure conditions, the target degradation order strategy (TDOS) and the trajectory online planning method were studied in this paper. Firstly, the influence of power failure on the orbit-injection process was analyzed. Secondly, the TDOS was proposed according to the mission attribute, failure mode, and multi-payload combination. Then, an online planning method based on the adaptive target update iterative guidance method (ATU-IGM) and radial basis neural network (RBFNN) was proposed, where the ATU-IGM adopted the basic TDOS criterion for generating optimal orbit-injection samples and online guidance instructions, and the RBFNN was used for orbit-injection samples training and online generation of orbital missions. Finally, the comparative simulation analysis was performed under multi-failure conditions. The results showed that the TDOS proposed in this paper could meet the requirements of the mission decision making under different failures, target orbit types, orbit-injection methods, and payload compositions. The online trajectory-planning capability deviation was less than 5%, and the mission decision-making and trajectory-planning time were less than 10 ms. This study provides theoretical support for autonomous decision making and planning of space launch missions.

Three-dimensional CFD investigation of a scramjet inlet under different freestream conditions

Scramjet (supersonic combustion ramjet) engines are often seen as a promising alternative to place payloads in the Earth's orbit. Such air-breathing engines have a simple structure and few moving parts. On the other hand, high heat fluxes and pressure loads on the walls, shock-wave-boundary-layer interactions, and the risk of choked flow inside the isolator channel are a few examples of obstacles that need to be addressed during scramjet design. Therefore, this study aims to evaluate of different air freestream conditions on the flow field in a scramjet inlet (compression ramps and isolator) through detailed three-dimensional computational fluid dynamics (CFD) simulations. The Mach number, flight altitude, and angle of attack were the evaluated conditions. Moreover, this study also focuses on the behavior of the boundary layer separation located at the compression ramp corners, isolator entrance, and scramjet sidewalls. Regarding the Mach number variation, the results showed that high Mach numbers yielded high-pressure levels throughout the engine. Furthermore, a higher altitude promoted a lower total pressure field. Considering the angle of attack changes, it is evident that a higher angle of attack results in a decrease in airflow pressure, while an increase in the total pressure along the walls is observed. By investigating these freestream parameters, this work can contribute to the early phase of engine design, avoiding critical failures in the scramjet structure due to aerodynamic load and thermal stress.

Measurement of Low-energy Cosmic-Ray Electron and Positron Spectra at 1 au with the AESOP-Lite Spectrometer

Abstract We report on a new measurement of the cosmic ray (CR) electron and positron spectra in the energy range of 20 MeV–1 GeV. The data were taken during the first flight of the balloon-borne spectrometer Anti Electron Sub Orbital Payload (AESOP-Lite), which was flown from Esrange, Sweden, to Ellesmere Island, Canada, in 2018 May. The instrument accumulated over 130 hr of exposure at an average altitude of 3 g cm−2 of residual atmosphere. The experiment uses a gas Cerenkov detector and a magnetic spectrometer, consisting of a permanent dipole magnet and silicon strip detectors (SSDs), to identify particle type and measure the rigidity. Electrons and positrons were detected against a background of protons and atmospheric secondary particles. The primary CR spectra of electrons and positrons, as well as the re-entrant albedo fluxes, were extracted between 20 MeV and 1 GeV during a positive solar magnetic polarity epoch. The positron fraction below 100 MeV appears flat, suggesting diffusion-dominated solar modulation at low rigidity. The all-electron spectrum is presented and compared with models from a heliospheric numerical transport code.

The Solar Orbiter mission

Aims.Solar Orbiter, the first mission of ESA’s Cosmic Vision 2015–2025 programme and a mission of international collaboration between ESA and NASA, will explore the Sun and heliosphere from close up and out of the ecliptic plane. It was launched on 10 February 2020 04:03 UTC from Cape Canaveral and aims to address key questions of solar and heliospheric physics pertaining to how the Sun creates and controls the Heliosphere, and why solar activity changes with time. To answer these, the mission carries six remote-sensing instruments to observe the Sun and the solar corona, and four in-situ instruments to measure the solar wind, energetic particles, and electromagnetic fields. In this paper, we describe the science objectives of the mission, and how these will be addressed by the joint observations of the instruments onboard.Methods.The paper first summarises the mission-level science objectives, followed by an overview of the spacecraft and payload. We report the observables and performance figures of each instrument, as well as the trajectory design. This is followed by a summary of the science operations concept. The paper concludes with a more detailed description of the science objectives.Results.Solar Orbiter will combine in-situ measurements in the heliosphere with high-resolution remote-sensing observations of the Sun to address fundamental questions of solar and heliospheric physics. The performance of the Solar Orbiter payload meets the requirements derived from the mission’s science objectives. Its science return will be augmented further by coordinated observations with other space missions and ground-based observatories.

The Spectrometer/Telescope for Imaging X-rays (STIX)

Aims. The Spectrometer Telescope for Imaging X-rays (STIX) on Solar Orbiter is a hard X-ray imaging spectrometer, which covers the energy range from 4 to 150 keV. STIX observes hard X-ray bremsstrahlung emissions from solar flares and therefore provides diagnostics of the hottest (⪆10 MK) flare plasma while quantifying the location, spectrum, and energy content of flare-accelerated nonthermal electrons. Methods. To accomplish this, STIX applies an indirect bigrid Fourier imaging technique using a set of tungsten grids (at pitches from 0.038 to 1 mm) in front of 32 coarsely pixelated CdTe detectors to provide information on angular scales from 7 to 180 arcsec with 1 keV energy resolution (at 6 keV). The imaging concept of STIX has intrinsically low telemetry and it is therefore well-suited to the limited resources available to the Solar Orbiter payload. To further reduce the downlinked data volume, STIX data are binned on board into 32 selectable energy bins and dynamically-adjusted time bins with a typical duration of 1 s during flares. Results. Through hard X-ray diagnostics, STIX provides critical information for understanding the acceleration of electrons at the Sun and their transport into interplanetary space and for determining the magnetic connection of Solar Orbiter back to the Sun. In this way, STIX serves to link Solar Orbiter’s remote and in-situ measurements.

A Space Inertial Sensor Ground Evaluation System for Non-Sensitive Axis Based on Torsion Pendulum

The inertial sensor is the key measurement payload of the technology verification satellite of China’s space gravitational wave detection mission-Taiji Project, which uses capacitive sensors to sense the acceleration disturbance of the test mass under the influence of non-conservative forces in the frequency range of 10 mHz~1 Hz. It is necessary to perform a ground performance evaluation and estimate the working state of the payload in orbit. However, due to the influence of the earth’s gravity and seismic noise, it is impossible to directly evaluate the resolution level of the non-sensitive axis when testing with high-voltage levitation, which leads to incomplete evaluation of the performance of the inertial sensor. In order to implement this part of the test, the sensitive structure is designed and a torsion pendulum facility for performance testing is developed. The experimental results show that the measurement resolution of the non-sensitive axis of the inertial sensor can reach 9.5 × 10−7 m/s2/Hz1/2 under the existing ground environmental conditions and is mainly influenced by the seismic noise during the system measurement. If the inertial sensor enters orbit, the measurement resolution can achieve 3.96 × 10−9 m/s2/Hz1/2, which meets the requirements of the technology verification satellite for a non-sensitive axis. This proposed system also provides a reasonable method for the comprehensive evaluation of inertial sensors in the future.

Exposure of structural epoxy adhesive to combination of tensile stress and γ-radiation

Mass is money! This sentence applies even more in space transportation than in aviation, due to the enormous cost it takes to carry payload in orbit. With the development of high temperature adhesives and the rise of composite materials like CFRP, adhesive bonding has experienced enormous growth in the space sector. Opposing to terrestrial application, the adhesive joint has to withstand unique environmental conditions. Adhesive systems belong to the group of polymers. Thus, the effect of radiation, especially ionizing radiation, has an enormous influence on the structural integrity of adhesive systems. In a space exploration mission e.g. on the lunar surface, the adhesive has to withstand mechanical loads in combination with radiation environment. This paper investigates the combined effect of tensile stress and γ-radiation on the mechanical properties of a 2-component epoxy adhesive. It's a matter of the greatest importance to investigate the possible degradation of the adhesive in combination with irradiation and tensile loading due to increased chain scission of stressed molecules. For structural adhesives, such phenomena are not covered well in the literature.In the present work, bulk specimens of the investigated adhesive are milled out from a plate and exposed to a60Co source. During the irradiation they were loaded in tensile direction. Afterwards, the specimens were inspected by FTIR spectroscopy and then tested destructively in order to determine the potentially degraded mechanical properties.The results show that the mechanical properties of the bulk adhesive stay constant for non-irradiated, irradiated non-loaded and irradiated loaded specimen which concludes that the investigated adhesive is not weakened by the combination of γ-radiation (32.4 kGy) and tensile loading.

Использование солнечной батареи в составе системы электропитания космического разгонного блока

A storage battery is routinely used in the existing boosters equipped with liquid-fuel rocket engines. Under certain conditions, the electric power supply system mass can be essentially decreased by using a storage battery in combination with a photovoltaic array. The article considers the possibilities of using, in the composition of the space booster (SB) power supply system (PSS), a photovoltaic array (PVA) in addition to a storage battery with the aim to decrease the PSS mass. A procedure for selecting the optimal storage battery and photovoltaic array characteristics ensuring the minimal PSS mass has been determined, and the values of these characteristics have been estimated taking as an example an SB intended for placing a payload in a geostationary orbit. The obtained results may find use in substantiating the characteristics and developing PSS elements for prospective boosters.

Boron based fillers as char enhancers of EPDM based heat shielding materials for SRMs: A comparative analysis

Solid Rocket Motors (SRMs) are used to launch satellites and other civilian or military payloads in orbit. Due to their low density and outstanding insulation capability, EPDM/aramid fibers ablatives represent the state of the art of Elastomeric Heat Shielding Materials (EHSMs) SRMs. However, this class of polymeric ablatives tend to exhibit a limited ablation resistance induced by thermo-oxidation and mechanical erosion due to the propellant combustion byproducts. In order to increase the thermo-oxidation resistance, char retention, and insulation index, boron based fillers can be considered in the material formulation of the SRM insulation liner. In this work we systematically studied the effect of different boron containing fillers embedded in EPDM based EHSMs. The heat capacity, thermal stability, mechanical properties, ablation resistance, pre- and post-oxy-acetylene torch testing morphology of these materials were analyzed evidencing the clear benefits of these fillers.

Improving of technical characteristics of launch vehicles with liquid rocket engines using active onboard de-orbiting systems

In this paper, the analysis of technical requirements (TR) for the development of modern space launch vehicles (LV) with main liquid rocket engines (LRE) is fulfilled in relation to the anthropogenic impact decreasing. Factual technical characteristics on the example of a promising type of rocket “Soyuz-2.1.v.” are analyzed. Meeting the TR in relation to anthropogenic impact decrease based on the conventional design approach and the content of the onboard system does not prove to be efficient and leads to depreciation of the initial technical characteristics obtained at the first design stage if these requirements are not included. In this concern, it is shown that the implementation of additional active onboard de-orbiting system (AODS) of worked-off stages (WS) into the onboard LV stages systems allows to meet the TR related to the LV environmental characteristics, including fire-explosion safety. In some cases, the orbital payload mass increases.

Human factors in launch flight safety

When developing a new launcher, the main design driver is operational efficiency (i.e. performance, reliability and availability), and cost effectiveness, not safety. Moreover, launch operators focus mission design on successfully placing the payload in orbit and on cost effectiveness. Those conditions are insufficient by themselves to ensure flight safety and may direct design decisions toward solutions that are less safe. The safety challenge of rocket launches typically requires additional dedicated safety equipment on-board and ground facilities to monitor flight for potentially unsafe conditions and to terminate the flight before the rocket can threaten population centers or critical assets. Flight Safety Officers have the ultimate responsibility to terminate the flight of an errant launcher before it endangers people and properties, by sending a command from ground. Flight Safety Officers must assess the situation and take action to terminate flight in a very short time, therefore human factors considerations are fundamental for correct and timely performance of the job. After a general overview of launch safety principles and practices, the paper illustrates the application of human factors considerations for flight safety officers with a discussion of training at the European spaceport in French Guiana.

Mechanical Limitations of Motorised Tethers for Payload Orbital Transfer

In recent decades much research has focused on the use of space tether systems for payload orbital transfers. Orbit raising through the exchange of momentum using hanging tether systems in addition to enhanced momentum exchange resulting from motorised torque-driven systems have both been considered. In this paper the fo- cus is on the performance of symmetrically laden motorised momentum exchange tethers with the aim of conducting orbital transfers. These consist of two propulsion tethers symmetrically attached to a motor shaft with payloads, to be later trans- ferred, attached to their free ends. Rotation is achieved through reaction against two further tethers and masses attached to the motor stator with an estimated mass of this system, minus the propulsion tethers and payloads, of 1.5 tonnes. By deriv- ing the maximum tension acting on a system orbiting an oblate Earth expressions are obtained for the maximum rotational velocity that the tether systems are capa- ble of withstanding, based upon the material properties of the tethers themselves. Having obtained these expressions the performance of both hanging and motorised tether systems for conducting orbital transfers is analysed in terms of the gain in energy when the orbital geometry and tether length are varied. From this analysis the material-specifc tether length which gives the greatest gain in rotational kinetic energy to the payloads is established for a motorised system. Finally, by utilising this material-specific length, the minimum semi-major axis which is capable of per- forming lunar transfers and Earth escape trajectories is obtained.

Water injection pre-compressor cooling assist space access

AbstractAdvances in space activity are linked to reductions in launch cost. Air-breathing propulsion-assisted flight systems offer the potential for revolutionary change of the space operations paradigm. Horizontal launch of a space-access system provides mission flexibility, responsiveness, and affordability. One way to reduce launch cost is to increase the Mach number at which a launch vehicle is staged from a carrier aircraft. Without exceeding the engine and airframe design limits, the pre-compressor cooling technology allows an operational aircraft to operate at Mach numbers and altitudes beyond its basic operational limits. This is an essential, near-term technology for reducing launch cost to place small-weight payloads in low Earth orbit. The advantage of this technology is assessed with a modified McDonnell Douglas QF-4C aircraft. Payloads are unachievable or marginal with an unmodified QF-4C. However, payloads weighing around 150 pounds are plausible with this aircraft when incorporating the water injection pre-compressor cooling (WIPCC) technology.

Scientific data products and the data pre-processing subsystem of the Chang'e-3 mission

The Chang'e-3 (CE-3) mission is China's first exploration mission on the surface of the Moon that uses a lander and a rover. Eight instruments that form the scientific payloads have the following objectives: (1) investigate the morphological features and geological structures at the landing site; (2) integrated in-situ analysis of minerals and chemical compositions; (3) integrated exploration of the structure of the lunar interior; (4) exploration of the lunar-terrestrial space environment, lunar surface environment and acquire Moon-based ultraviolet astronomical observations. The Ground Research and Application System (GRAS) is in charge of data acquisition and pre-processing, management of the payload in orbit, and managing the data products and their applications. The Data Pre-processing Subsystem (DPS) is a part of GRAS. The task of DPS is the pre-processing of raw data from the eight instruments that are part of CE-3, including channel processing, unpacking, package sorting, calibration and correction, identification of geographical location, calculation of probe azimuth angle, probe zenith angle, solar azimuth angle, and solar zenith angle and so on, and conducting quality checks. These processes produce Level 0, Level 1 and Level 2 data. The computing platform of this subsystem is comprised of a high-performance computing cluster, including a real-time subsystem used for processing Level 0 data and a post-time subsystem for generating Level 1 and Level 2 data. This paper describes the CE-3 data pre-processing method, the data pre-processing subsystem, data classification, data validity and data products that are used for scientific studies.