Satellite Launch Vehicles

During the first two decades of the 21st century, the interest in the Moon missions' realization has been updated from the side of states and commercial operators. Many governmental, commercial and academic Moon's missions are under development now. The development of the information system for researching the Moon's surface and supporting the Lunar human and unmanned missions is actual. The Conceptual approach in the development of the Selena-Information System is presented. The Selena-Information system is intended to provide the Earth's and Moon's consumers with multi-layer information about the Moon's surface and distribution of the soils including minerals and water. The Selena-Information System combines three parts: the Moon segment, the Earth segment, and the Telecommunication segment. The raw Moon remote sensing information could be obtained by the constellation of small satellites for remote sensing of the Moon in the low Moon Orbit. Taking into account the low power capacity of the small satellite the huge information volume can be transmitted in a short distance. It is proposed to put the two Lunar Landing Modules in the Moon's polar regions. These Modules provide control of the small satellite in the Low Moon Orbit, receiving raw data, processing, and storage of it. The Earth segment includes the system control mission center, Data Centers of the Selena-Information System Information consumers, control mission centers of consumer's systems. The telecommunication segment provides the connection between the Moon and Earth segments and uses Big Antennas and ground communication systems.

A design procedure to realize a hybrid rocket engine that uses hydrogen peroxide as the oxidizer and a solid polymer as the fuel is presented. In the design procedure, the products of decomposition and of combustion at different stations are calculated by linking a standard complex chemical-equilibrium program of NASA. For the designed basic-configuration, applying the mass conservation principle, theoretical performance results are presented with propellant-tank pressure and grain port length as parameters.

This chapter briefly introduces hybrid rocket propulsion for general audience. Advantageous of hybrid rockets over solids and liquids are presented. This chapter also explains how to design a test setup for hybrid motor firings. Hybrid propulsion provides sustainable, safe and low cost systems for space missions. Therefore, this chapter proposes hybrid propulsion system for Mars Ascent Vehicles. Paraffin wax is the fuel of the rocket. Propulsion system uses CO 2 =N 2 O mixture as the oxidizer. The goal is to understand the ignition capability of the CO 2 as an in-situ oxidizer on Mars. CO 2 is known as major combustion product in the nature. However, it can only burn with metallic powders. Thus, metallic additives are added in the fuel grain. Results show that CO 2 increase slows down the chemical kinetics thus reduces the adiabatic flame temperature. Maximum flammability limit is achieved at 75% CO 2 by mass in the oxidizer mixture. Flame temperature is 1700 K at 75% CO 2 . Ignition quenches below the 1700 K.

The problem of an affordable, responsive, and reliable microsat launch system (MLS) has bedeviled the small-satellite community, especially in the United States, for decades. Rides on dedicated vehicles may cost $15M or more, while shared space is hard to find and harder to fit with the schedules of microsat operators. Several efforts to build an MLS, generally focusing on cheap expendable launch vehicles (ELVs), have failed, as did the government's much-touted Bantam launch vehicle effort in the 1990s. Today, there are several options, both reusable and expendable, in development, as government agencies and corporations respond to the growing interest in microsats by trying once again to solve the problem. In pursuing these efforts, it is instructive to consider why the problem was not solved long ago. A reliable and relatively affordable MLS, the NASA-developed Scout, was built over four decades ago. Since then, technological advances should have made duplicating its success a relatively minor problem. Why has this not been so? The answers range from the volatile microsat launch market to the fixed costs involved in launch ranges and safety standards, to the technology itself. This paper examines MLS development efforts past and present, analyzes the technical and economic factors retarding their success, and offers prescriptions for the organizations now attacking the MLS problem. American Institute of Aeronautics and Astronautics why the problem has resisted solution. This paper examines MLS development efforts past and present, analyzes the technical, economic, and organizational factors retarding their success, and offers prescriptions for those now engaged in the MLS field. This paper focuses on the most difficult part of the microsat launch problem -development of a dedicated U.S. vehicle that will lower costs sufficiently to bring about expanded use of microsats and remove launch as a constraint.

Affordable, reliable, and responsive launch is critical to the survival and growth of small satellite systems. Unfortunately, no small launch vehicle now on the U.S. market offers a perfect solution: low cost, proven reliability, and the flexibility to meet all user requirements. There are several small launchers on the market, and choosing the best launcher for a given mission is a difficult challenge for the smallsat developer. Using the launch of three sample smallsats as a mission model, this paper surveys launch vehicle options currently available or in advanced development. Those launch vehicles capable of launching the mission model satellites into the desired orbits are then evaluated on cost, risk, availability, and payload environment. A combined ranking produced by a computer model indicates the optimal choice for each mission. This exercise demonstrates that selection of the best option does not always mean picking the obvious or best-known vehicle, and that small satell...

The results from calculating the apogee altitude for different 1u configuration payloads show that the maximum altitude expected is 21,000 ft. for a 1U payload and 11,000 ft. for a 6U payload. The motor chosen for the launch is the Pro 150 O5800 White Thunder with a total impulse of 30,601 Ns, average thrust of 5,802 N, and burn time of 5.27 seconds. OpenRocket software was used to simulate the flight profile and the loads expected during the launch. Previous Work

For over a decade, student ballooning projects have provided a wonderful opportunity of launching small scientific payloads to near space environment in a small budget. Apart from the obvious merits of such projects, one significant outcome is the cross disciplinary training for undergraduate science and engineering students that prepares them for a future career in industry. Students are also exposed to research technique that is a strong motivational factor to work towards a graduate degree. A faculty/student team from Albany State University was one of the two teams that participated in one such NSF sponsored ballooning program in the summer of 2009 conducted by Louisiana State University. The teams were provided training to design, develop and fabricate the payloads that were launched from NASA's Columbia Scientific Balloon Facility at Palestine, Texas followed by data retrieval and science presentations. The summer program is being followed by a yearlong activity of designing and building a payload for a new experiment at the home institution that will be launched in late spring.

For over a decade, student ballooning projects have provided a wonderful opportunity of launching small scientific payloads to near space environment in a small budget. Apart from the obvious merits of such projects, one significant outcome is the cross disciplinary training for undergraduate science and engineering students that prepares them for a future career in industry. Students are also exposed to research technique that is a strong motivational factor to work towards a graduate degree. A faculty/student team from Albany State University was one of the two teams that participated in one such NSF sponsored ballooning program in the summer of 2009 conducted by Louisiana State University. The teams were provided training to design, develop and fabricate the payloads that were launched from NASA's Columbia Scientific Balloon Facility at Palestine, Texas followed by data retrieval and science presentations. The summer program is being followed by a yearlong activity of designi...

Increased performance with reduced overall cost, and precise design and optimization of launch systems are critical to affordability. In this respect, the use of hybrid motors has increased to ease handling based on motor selection. In the current study, the launch vehicle's performance is enhanced by incorporating a hybrid rocket motor into the last stage and optimized using particle swarm optimization to develop a six-degrees-of-freedom tool. This modification aims to increase payload placement flexibility, facilitate handling, and reduce costs. Thanks to the interactive subsystems within this research, this innovative study more comprehensively considers the launch vehicle trajectory design problem, allowing the simultaneous consideration of the effect of launch vehicle geometry along with other parameters in the system. In this context, the proposed method is applied to the Minotaur-I launch vehicle, and contributions of the detailed design and optimization are presented. Optimization results show that the percentage differences between these models for the original vehicle were observed to be 11.55% in velocity and 8.02% in altitude. However, there were differences of 10.06% and 48.8%, 15.8% and 23.2%, and 19.5% and 78.9% in altitudes and velocities when the center of gravity and moment of inertia changes were neglected, and constant aerodynamic coefficients were assumed, respectively. In all these cases, it was observed that the flight path angle was not close to zero. Moreover, the same mission was achieved by the launch vehicle with the optimized hybrid last stage and the propulsion performance was increased by about 7.64% based on the specific impulse and total impulse-overweight ratio.

Optimization of Hybrid Rocket Engines at Politecnico di Torino began in the 1990s. A comprehensive review of the related research activities carried out in the last three decades is here presented. After a brief introduction that retraces driving motivations and the most significant steps of the research path, the more relevant aspects of analysis, modeling and achieved results are illustrated. First, criteria for the propulsion system preliminary design choices (namely the propellant combination, the feed system and the grain design) are summarized and the engine modeling is presented. Then, the authors describe the in-house tools that have been developed and used for coupled trajectory and propulsion system design optimization. Both deterministic and robust-based approaches are presented. The applications that the authors analyzed over the years, starting from simpler hybrid powered sounding rocket to more complex multi-stage launchers, are then presented. Finally, authors’ conclu...

An electrically driven pump-fed cycle for a hybrid rocket engine is proposed and compared to a simpler gas-pressurized feed system. A liquid-oxygen/paraffin-based fuel hybrid rocket engine which powers the third stage of a Vega-like launcher is considered. Third-stage ignition conditions are assigned, and engine design and payload mass are defined by a proper set of parameters. Uncertainties in the classical regression rate correlation coefficients are taken into account and robust design optimization is carried out with an approach based on an epsilon-constrained evolutionary algorithm. A mission-specific objective function, which takes into account both the payload mass and the ability of the rocket to reach the required final orbit despite uncertainties, is determined by an indirect trajectory optimization approach. The target orbit is a 700 km altitude polar orbit. Results show that electrically driven pump-fed cycle is a viable option for the replacement of the conventional gas...

In recent years CubeSats have transcended their original purpose of being an educational and technology demonstration platform only. It has now brought in reach affordable yet state-of-the-art Earth observation missions to academia, governments, and private industry. ISISpace has taken up its share by supporting this Earth observation trend with the launch of NAPA-1 in September 2020, it being the first Earth Observation CubeSat of the Royal Thai Airforce. Successful commissioning and the transfer of operations to the customer has led to valuable feedback and lessons learnt. These have been directly put to use during the design and development of its successor NAPA-2. NAPA-2 takes small satellite Earth observation up a notch thanks to its multispectral imager: the Simera MultiScape100 CIS. This not only means an upgrade from NAPA-1's 50-meter ground sampling distance (GSD) to a stunning 5-meter GSD, but also being able to scan the surface in up to 7 spectral bands at the same time. Implications of this, however, are stricter platform requirements to meet optical performance and the need for dedicated on-ground image processing capability. The latter is achieved by the inclusion of Pinkmatter Solutions' FarEarth application. While images and satellite data obtained and lessons learnt throughout the NAPA-1 commissioning period will be shared, this paper will focus on the improvements made and the resulting differences with the next generation satellite NAPA-2. Insight into its design and in-orbit performance will be provided. On top of that, the assembly, integration, and verification phase will be elaborated on. It will also include details on the on ground optical verification of the MultiScape100 CIS, showing how verification before and after, for example, environmental testing was achieved. Finally, an outlook will be given to show the next steps in the coming years at ISISpace regarding small satellite Earth observation and image processing.

The launch phase of a suborbital rocket is always associated with severe structural stresses and they need to be cautiously accounted for during the design level. The structural testing and the numerical analysis of six degrees of freedom and variable mass multi-stage suborbital launcher at launch are reported here. The paper describes the results of free-free modal analysis followed by static and dynamic finite element model analyses of a three-stage rocket prototype subject to launch stresses. The numerical natural frequencies are compared with an experimental modal analysis conducted with the Impact Hammer Modal Testing method. The dynamic results obtained with a finite element model are also compared with dynamic tests in order to validate the numerical model. These results are part of the ROSA-STAR national project "Technologies for Testing and Validation of the Structure, and Modified Hybrid Rocket Motor for Suborbital Launcher-STRAC".

, (650)691-0676 As the utility and usefulness of small spacecraft are being demonstrated in orbit by NASA and other organizations and academia, it is becoming clear that small spacecraft can and should have a large roll in the exploration of the solar system. The various benefits associated with small platforms include low cost, frequent missions, and by virtue of those features, repeatable and evolutionary missions. This paper describes the motivation for establishing a dedicated function within the NASA Ames Research Center to pursue and implement multiple missions with small, inexpensive spacecraft, and outlines some future trends and directions of the Small Spacecraft Division within the context of science and exploration.

Climate change poses a significant threat to our planet, and accurate monitoring of essential climate variables is crucial for understanding and mitigating its effects. However, current remote sensing missions do not provide systemic high quality spatial resolution data for large swath areas. It is also unclear as to whether these missions are sustainable and follow the recommended international guidelines for space missions. The main objective of this research paper is to conceptualize a climate change monitoring satellite mission that is proven to be sustainable through life cycle assessment (LCA). LCA is an important tool that can help evaluate conceptual designs based on their environmental impact over their life cycles and when employed in the space mission design context, it can help determine efficient and sustainable alternatives to presently used materials and processes in satellites. This research focuses on the conceptual design of a constellation of low altitude remote sensing small satellites at SSO (Sun Synchronous Orbit) and orbits having a medium range inclination-MIO (Mid Inclination Orbit), in order to obtain frequent coverage at both higher and lower altitudes with greater revisit frequency. These satellites monitor key climate change indicators as defined by the EPA, such as surface temperature, polar ice cap melting rates, percentage of sea level rise, precipitation rates and greenhouse gas emissions among many more through novel instruments that are designed following the heritage of missions like Copernicus-Sentinel, LandSat and A-Train and improving upon the resolution obtained using recently introduced technology like hyperspectral imagery and high speed onboard data processing systems. The research also includes the design parameters of the satellite's structure, electronics and power handling system, orbit determination and attitude control system, onboard electric propulsion system and data handling and communication system. The proposed mission design also complies with the existing space law policies and has proper deorbit procedures like the recently introduced 5 year deorbit rule adopted by the FCC. In summary, this paper outlines the conceptual design of a sustainable climate change monitoring satellite mission that provides high-resolution, frequent coverage of key climate indicators over large areas, thus providing a solution to the gap present in the existing remote sensing satellite missions.

This work presents a multidisciplinary optimization technique applied to the design of a hybrid rocket motor for the de-boost system of a small recoverable satellite. For some specific missions, hybrid rocket technology may become a competitive choice due to its inherent advantages over more traditional chemical propulsion systems. The Brazilian Recoverable Satellite Platform (SARA) relies on a chemical propulsion system for orbital de-boost. We considered a hybrid motor as the propulsive technology. An optimization algorithm was developed to help propose a motor configuration. We choose a hybrid motor based on liquefying fuel, solid paraffin, and nitrous oxide propellants, by virtue of the long term tradition in operating such propulsion systems. The multidisciplinary configuration optimization technique was entirely based on geometrical operating parameters of the motor, rather than performance to facilitate further design and fabrication. Results from the code presented a hybrid ...

The SARA satellite was conceived as a microgravity recoverable and reusable research platform by both Brazilian Institute of Aeronautics and Space (IAE) and Space Agency (AEB). Figure 1a shows an artistic design of the platform and Fig. 1b, the planned launch vehicle of the spacecraft. The satellite was designed to carry a 55 kg payload mass, with total launch mass not exceeding 350 kg. Missions were projected to take place in circular orbits at 300 km altitude, with two-degree inclination. After completion of the microgravity experiments, up to ten days, the reentry procedure began by providing the right positioning of the satellite followed by the deboost impulse. Final deceleration took place by a high performance parachute system (Koldaev and Moraes, 1997). The spacecraft was scheduled to pass through a series of uali cation tests in ballistic ights reaching 350 km apogee and falling at about 300 km from the launch site. The system should be accelerated by a Brazilian VS-40 sounding rocket (Fig. 1b). To date, solid and liquid rocket propulsion systems were the only technological means considered for the deboost

The launch phase of a suborbital rocket is always associated with severe structural stresses and they need to be cautiously accounted for during the design level. The structural testing and the numerical analysis of six degrees of freedom and variable mass multi-stage suborbital launcher at launch are reported here. The paper describes the results of free-free modal analysis followed by static and dynamic finite element model analyses of a three-stage rocket prototype subject to launch stresses. The numerical natural frequencies are compared with an experimental modal analysis conducted with the Impact Hammer Modal Testing method. The dynamic results obtained with a finite element model are also compared with dynamic tests in order to validate the numerical model. These results are part of the ROSA-STAR national project "Technologies for Testing and Validation of the Structure, and Modified Hybrid Rocket Motor for Suborbital Launcher-STRAC".

In 2023, satellite communications experienced unprecedented advancements in both operational efficiency and satellite launch capabilities, setting the stage for a dynamic trajectory. Experts in the satellite communications sector contend that there will be more advances in 2024, such as the integration of multiple orbits spanning high-, medium-, and low-Earth orbits. Learn more in Satellite Communication Training courses by Tonex.

In the Frame of Romanian Incentive Scheme Programme under European Space Agency, a feasibility study of a Small Orbital Launcher was performed by "Politehnica" University of Bucharest. Building on the results obtained in the project, European Space Agency's Space Transportation Directorate through the Future Launchers Preparatory Programme awarded INCAS a Phase 0/A contract for a Microlauncher concept. The targeted market is the micro-satellites sector developed by universities, but also by small companies for research purpose and validation of applications/products before commercialization, with a mass lower than 150 kg. This paper presents a conceptual design for a Microlauncher having a target mission to deliver a small satellite of 150 kg into a 600 km circular polar orbit with the launching site located in Europe. The Microlauncher concept is obtained using an in-house multidisciplinary optimization tool, which contains four modules: weights and sizing, aerodynamics performance assessment, propulsion, trajectory computation and optimization.

Dr. Frank J. Malina, (1912-81) is best known to members of the International Astronautical Federation and the International Academy of Astronautics for his deep commitment to, and active involvement in, both organizations; to his associates at UNESCO for his leadership in organizing the UNESCO Division of Scientific Research; to members of the art world for his kinetic paintings and editorship of the journal Leonardo; and to aerospace historians for his co-founding, with Dr. Theodore yon Karman (1881-1963), of the Jet Propulsion Laboratory with its World War II development of rocket takeoff assist units for aircraft and, after WWll, America's first production sounding rocket, the WAC Corporal. However, Frank Malina has received little recognition from aerospace educators for his equally pioneering efforts on their behalf. This paper highlights Dr. Malina's efforts not only as an early participant in educatorsupported student rocketry but also in helping establishing some of the guiding principles of this field of science motivation. As Dr. Malina noted in 1968, upon becoming one of the founding members of the Supervision of Youth Research Experiments (SYRE) subcommittee of the 1AF's Education Committee, the fundamental safety and educational ground rules of: (A) qualified supervision, (B) proper safety facilities, and (C) professionally designed equipment, conceived almost fifty years ago, are equally valid to today's supervised youth rocketry space-related experiment educational programs.

An electrically driven pump-fed cycle for a hybrid rocket engine is proposed and compared to a simpler gas-pressurized feed system. A liquid-oxygen/paraffin-based fuel hybrid rocket engine which powers the third stage of a Vega-like launcher is considered. Third-stage ignition conditions are assigned, and engine design and payload mass are defined by a proper set of parameters. Uncertainties in the classical regression rate correlation coefficients are taken into account and robust design optimization is carried out with an approach based on an epsilon-constrained evolutionary algorithm. A mission-specific objective function, which takes into account both the payload mass and the ability of the rocket to reach the required final orbit despite uncertainties, is determined by an indirect trajectory optimization approach. The target orbit is a 700 km altitude polar orbit. Results show that electrically driven pump-fed cycle is a viable option for the replacement of the conventional gas-pressurized feed system. Robustness in the design is granted and a remarkable payload gain is achieved, using both present and advanced technologies for electrical systems.

To settle the question of the nature of the interstellar object 1I/'Oumuamua requires in-situ observations via a spacecraft, as the object is already out of range of existing telescopes. Most previous proposals for reaching 1I/'Oumuamua using near-term technologies are based on the Solar Oberth Manoeuvre (SOM), as trajectories without the SOM are generally significantly inferior in terms of lower mission duration and higher total velocity requirement. While the SOM allows huge velocity gains, it is also technically challenging and thereby increases programmatic and mission-related risks. In this paper, we identify an alternative route to the interstellar object 1I/'Oumuamua, based on a launch in 2028, which does not require a SOM but has a similar performance as missions with a SOM. It instead employs a Jupiter Oberth Manoeuvre (JOM) with a total time of flight of around 26 years or so. The efficacy of this trajectory is a result of it significantly reducing the ∆V to Jupiter by exploiting the VEEGA sequence. The total ∆V of the trajectory is 15.8 km s −1 and the corresponding payload mass is 115 kg for a SLS Block 1B or 241 kg for a Block 2. A further advantage of the JOM is that the arrival speed relative to 1I/'Oumuamua is approximately 18 km s −1 , much lower than the equivalent for the SOM of around 30 km s −1 .

A tube-in-tube concept for separation of bodies in space was investigated theoretically and experimentally. The separation system is based on generation of high pressure gas by combustion of solid propellant and restricting the expansion of the gas only by ejecting the two bodies in opposite directions, in such a fashion that maximizes generated impulse. An interior ballistics model was developed in order to investigate the potential benefits of the separation system for a large range of space body masses and for different design parameters such as geometry and propellant. The model takes into account solid propellant combustion, heat losses, and gas phase chemical reactions. The model shows that for large bodies (above 100 kg) and typical separation velocities of 5 m/s, the proposed separation mechanism may be characterized by a specific impulse of 25,000 s, two order of magnitude larger than that of conventional solid rockets. It means that the proposed separation system requires only 1% of the propellant mass that would be needed for a conventional rocket for the same mission. Since many existing launch vehicles obtain such separation velocities by using conventional solid rocket motors (retro-rockets), the implementation of the new separation system design can reduce dramatically the mass of the separation system and increase safety. A dedicated experimental setup was built in order to demonstrate the concept and validate the model. The experimental results revealed specific impulse values of up to 27,000 s and showed good correspondence with the model.

have been part of the Regional Revitalization Promotion Project by the Center of Community (and Intelligence) University (COC+) program of Gifu University's Regional Cooperation Center. This program has given high school students from the Gifu prefecture the opportunity to experience the education and research of Gifu University in a high school-university collaboration project using the video conference system provided by the Ministry of Education, Culture, Sports, Science and Technology. The project conveys the charm of the aerospace industry in the Gifu prefecture to foster and arouse motivation to work in the industry, provide basic knowledge about space engineering, and play an active part in the future artificial satellite, rocket, aviation industry. We thus held a space engineering course for the purpose of developing human resources.

The Passepartout sounding balloon transportation system for low-mass (<1200 g) experiments or hardware for validation to an altitude of 35 km is described. We present the general flight configuration, setup of the flight control system, environmental and position sensors, power system, buoyancy considerations as well as the ground control infrastructure including recovery operations. In the telemetry and command module the integrated airborne computer is able to control the experiment, transmit telemetry and environmental data and allows for a duplex communication to a control centre for tele-commanding. The experiment module is mounted below the telemetry and command module and can either work as a standalone system or be controlled by the airborne computer. This spacing between experiment-and control unit allows for a high flexibility in the experiment design. After a parachute landing, the on-board satellite based recovery subsystems allow for a rapid tracking and recovery of the telemetry and command module and the experiment. We discuss flight data and lessons learned from two representative flights with research payloads.

Pioneer missions used gravity assists to insert the spacecraft on solar system escape trajectories, whereas Mariner 10, Galileo, Cassini, and Messenger missions exploited inner-planetary gravity assists to reach Mercury, Jupiter, and Saturn, respectively, without the need for heavy-lift launch vehicles [10-12]. As missions become increasingly complex, determining feasible trajectories quickly becomes a daunting task. These types of interplanetary and interstellar mission design problems present significant optimization challenges. They are extremely nonlinear, often with multiple strong basins of attraction in the

The WinCube Satellite Project is a cooperative effort among Manitoba high schools, the Manitoba Satellite Interest group (MSIG), the Faculty of Engineering at the University of Manitoba, Maples Collegiate Space Exploration Academy, the Manitoba Aerospace Human Resources Coordinating Committee and numerous aerospace industry partners. Through a mentorship program, Manitoba high school students will design, construct, and launch a pico-satellite with technical support provided by aerospace faculty and engineering students. Basic system design and construction experience for the high school students is provided by the construction and launch of high altitude balloon payloads. Students learn first hand about space mission design, telecommunications, programming, electrical and mechanical engineering.

Bungo Shiotani is a Ph.D. student at the University of Florida (UF) working on systems engineering aspects for small satellites. Specifically to develop metrics to quantify mission assurance throughout the project life-cycle. Bungo received two Bachelor of Science degrees, one in Aerospace Engineering from UF and the other in Engineering Physics from Jacksonville University. He also received his Master of Science in Aerospace Engineering from UF where his thesis, Reliability Analysis of SwampSat, focused on performing reliability analyses on SwampSat, UF's first CubeSat. His experiences and as the project manager with SwampSat lead to an internship at NESTRA (Japan) where he worked on developing system diagrams and test procedures as well as assembly integration and testing of their three microsatellites that were in development. In addition to his Ph.D. work, Bungo is the project coordinator for Partnerships for International Research and Education (PIRE) program on multiphase fluid science and technologies at the UF's Chemical Engineering Department funded by the National Science Foundation. As the PIRE project coordinator, he works as a liaison with members from France, Japan, and the United States. Mr. Dante Augustus Buckley, UF's Space Systems Group Dante received his Bachelor of Science in Aerospace Engineering from the University of Florida (UF) in Gainesville in 2006. He is credited for co-founding the Small Satellite Design Club (SSDC) in his undergraduate years after being selected to compete in the Frank J. Redd student competition at the 19th Annual USU/AIAA Small Satellite Conference in Logan, Utah. SSDC is a UF affiliated student organization whose primary goal was to establish a permanent small satellite program on campus while creating awareness of space systems engineering in order to prepare students for their professional careers. Dante is a consultant (former research assistant) for the Space Systems Group (SSG), a graduate research team advised by Professor Norman Fitz-Coy in the Department of Mechanical & Aerospace Engineering. SSG designed and developed a cube satellite mission known as SwampSat, which launched in 2013. SwampSat is a CubeSat for on-orbit demonstration of a compact three-axis attitude control system developed at UF geared to affect rapid retargeting and precision pointing (R2P2) of pico-class (1 kg) and nano-class (<10 kg) spacecraft. Through Dante's leadership, SSDC won the Annual Florida University Nano-SATellite (FUNSAT) design competition sponsored by the Florida Space Grant Consortium and Space Florida in both 2008 and 2009. Dante is also credited for establishing and coordinating outreach efforts. He is the Lead Architect on the Educate Utilizing CubeSat Experience (EdUCE) project, a Science, Technology, Engineering, Arts and Mathematics (STEAM) and CubeSat themed K-20 outreach initiative, which in return just may inspire and cultivate the next generation of scientists and engineers to carry this and other work forward. Dante earned a Dean's Certificate in Engineering Entrepreneurship in Spring 2011 through the Engineering Leadership Institute. Dante is the former Curriculum Technologist / Computer Applications instructor at St. Francis Catholic High School (SFCHS) in Gainesville, FL. where he also hosted a TEDxYouth event and coached a FIRST robotics rookie all-star team. Dante is currently a Manufacturing Engineer II at RTI Surgical where he supports the manufacturing process for surgical implants.

The Moon is largely untouched by problems of space junk, but that could change as more companies and countries set their sights on it. © ESA/NASA On 4 March, a discarded rocket booster will slam into Hertzsprung Crater on the far side of the Moon-humanity's first known piece of litter to unintentionally reach the lunar surface. Skywatchers originally thought the booster was part of a SpaceX Falcon 9 rocket launched in 2015. But now they say it is part of a Chinese Long March 3C rocket that blasted off in 2014.

Nanosatellites are becoming a preferred platform for testing innovative technologies and conducting academic research in space. Flight Software (FSW) is a software program that operates on a processor embedded within the satellite structure (Miranda et al., 2019). It assumes the crucial responsibility of managing various aspects of satellite operations, including activity control, data processing, and ensuring satellite health and safety. By fulfilling these tasks, the FSW empowers the satellite to execute all actions essential for achieving its scientific objectives.

This paper proposes a new approach in multi-step sequential system optimization (MSSO) to implement a conceptual design for satellite upper stage with a maneuver in the conditions close to reality. In this method of design, there are two main cycles; trajectory optimization cycle and optimal design cycle, each one is correlated to each other in another cycle called configuration. In the trajectory optimization, the optimization objective is to place the upper stage in the destination orbit, using the minimum amount of fuel consumption. In this cycle, a new approach has been introduced for a three-dimensional trajectory using two genetic algorithms inside each other. In addition, selecting the suitable engine is carried out in this cycle. Convergence of design and exclusion of design are carried out in the configuration cycle. Convergence and optimization of subsystems design are carried out in the optimal design cycle. The innovations of this paper are implementation of the design according to multi-step sequential system design in which optimization is performed step by step, and orbital optimization is introduced according to a new approach. Choosing a desirable criterion for optimization process and proper coefficient for convergence in design, are among considerable characteristics of this paper. Validation has been performed using one of the upper stages in the world.

The Passepartout sounding balloon transportation system for low-mass (<1200 g) experiments or hardware for validation to an altitude of 35 km is described. We present the general flight configuration, setup of the flight control system, environmental and position sensors, power system, buoyancy considerations as well as the ground control infrastructure including recovery operations. In the telemetry and command module the integrated airborne computer is able to control the experiment, transmit telemetry and environmental data and allows for a duplex communication to a control centre for tele-commanding. The experiment module is mounted below the telemetry and command module and can either work as a standalone system or be controlled by the airborne computer. This spacing between experiment-and control unit allows for a high flexibility in the experiment design. After a parachute landing, the on-board satellite based recovery subsystems allow for a rapid tracking and recovery of the telemetry and command module and the experiment. We discuss flight data and lessons learned from two representative flights with research payloads.

Nanosatellites are becoming a preferred platform for testing innovative technologies and conducting academic research in space. Flight Software (FSW) is a software program that operates on a processor embedded within the satellite structure (Miranda et al., 2019). It assumes the crucial responsibility of managing various aspects of satellite operations, including activity control, data processing, and ensuring satellite health and safety. By fulfilling these tasks, the FSW empowers the satellite to execute all actions essential for achieving its scientific objectives.

The paper deals with the conceptual design of a space tug to be used in support to Earth satellites transfer manoeuvres. Usually Earth satellites are released in a non-definitive low orbit, depending on the adopted launcher, and they need to be equipped with an adequate propulsion system able to perform the transfer to their final operational location. In order to reduce the mass at launch of the satellite system, an element pre-deployed on orbit, i.e. the space tug, can be exploited to perform the transfer manoeuvres; this allows simplifying the propulsion requirements for the satellite, with a consequent decrease of mass and volume, in favour of larger payloads. The space tug here presented is conceived to be used for the transfer of a few satellites from low to high orbits, and vice versa, if needed. To support these manoeuvres, dedicated refuelling operations are envisaged. The paper starts from on overview of the mission scenario, the concept of operations and the related architecture elements. Then it focuses on the detailed definition of the space tug, from the requirements' assessment up to the budgets' development, through an iterative and recursive design process. The overall mission scenario has been derived from a set of trade-off analyses that have been performed to choose the mission architecture and operations that better satisfy stakeholder expectations: the most important features of these analyses and their results are described within the paper. Eventually, in the last part of the work main conclusions are drawn on the selected mission scenario and space tug and further utilizations of this innovative system in the frame of future space exploration are discussed. Specifically, an enhanced version of the space tug that has been described in the paper could be used to support on orbit assembly of large spacecraft for distant and long exploration missions. The Space Tug development is an activity carried on in the frame of the SA-PERE project (Space Advanced Project Excellence in Research and Enterprise), supported by Italian Ministry of Research and University (MIUR), and specifically in its STRONG sub-project (Systems Technology and Research National Global Operations) and related to the theme of space exploration and access to space. * 2000kg has been assumed as a preliminary estimation, further analyses will be carried to refine this assessment.

The VSB-30 vehicle is a two-stage, unguided, rail launched sounding rocket, consisting of two solid propellant motors, payload, with recovery and service system. By the end of 2010, ten vehicles had already been launched, three from Brazil (Alcântara) and seven from Sweden (Esrange). The objective of this paper is to give an overview of the main characteristics of the first ten flights of the VSB-30, with emphasis on performance and trajectory data. The circular 3σ dispersion area for payload impact point has around 50 km of radius. In most launchings of such vehicle, the impact of the payload fell within 2 sigma. This provides the possibility for further studies to decrease the area of dispersion from the impact point.

Cosmos Aerospace Association is a leading engineering students’ group, located in the Universidad Rey Juan Carlos (URJC) in Madrid, Spain. Providing a one-of-a-kind opportunity to all varieties of students for both personal and engineering growth, it is one of the few active aerospace student associations in Spain. Within this work, we introduce the achievements, influence and lessons learned from our association in these years. We focus on its educational impact in the environment of the university: not only from the perspective of aerospace-related degrees but also in the promotion of STEM careers on students of all ages. Conceived by undergraduate aerospace students and supported by professors and university staff, Cosmos was born to provide a creative and learning environment in the promotion of our passion for space and science in general. Bringing together students with similar mindsets, it has become a symbiotic platform in which all university actors share their efforts and ...

High-altitude balloons (HABs) present a valuable and cost-effective tool for educators and students to access the conditions that are analogous to space and extraterrestrial environments in the Earth’s upper atmosphere. Historically, HABs have been used for meteorological measurements, observation, sampling of aerosols, and exposure of samples to upper atmosphere environments. The Earth’s stratosphere allows researchers access to a unique combination of wideband solar radiation, extreme cold, rarefied air, low humidity, and acute ionizing radiation—conditions that are relevant to space biology research. Here, we describe a reproducible payload for a HAB mission that can be constructed, launched, and retrieved for about $3,000. This general standard operating procedure can be used by educators, community scientists, and research teams working with limited resources.

High-altitude balloons (HABs) present a valuable and cost-effective tool for educators and students to access the conditions that are analogous to space and extraterrestrial environments in the Earth’s upper atmosphere. Historically, HABs have been used for meteorological measurements, observation, sampling of aerosols, and exposure of samples to upper atmosphere environments. The Earth’s stratosphere allows researchers access to a unique combination of wideband solar radiation, extreme cold, rarefied air, low humidity, and acute ionizing radiation—conditions that are relevant to space biology research. Here, we describe a reproducible payload for a HAB mission that can be constructed, launched, and retrieved for about $3,000. This general standard operating procedure can be used by educators, community scientists, and research teams working with limited resources.

Reusability in Launchers has been a topic of research and interest in the last years to improve the Access to Space. In Europe, PLD Space is one of the few companies betting in the concept of recovery and reusability in the launchers sector. With MIURA1, a sounding rocket and technology demonstrator of a microlauncher, PLD Space wants to demonstrate a potential recovery and reusability strategy for small launchers. MIURA1, designed to be used up to three times, is a liquid propellant rocket with the capability of carrying up to 100 kg of payload to an apogee of 120 km. Its recovery system is characterized by a deceleration system and a two-stage parachutes system. The sounding rocket will be launched from CEDEA Launch Site at the South of Spain in 2020 and will be recovered after splashdown at around 80 km (nominal case) from the Gulf of Huelva. MIURA1, after reaching apogee, will slow-down until it will reach handover conditions for the two- stage parachutes opening. The first-stag...

The paper presents the influence of the main design requirements on the optimal two-stage small launcher configuration. The influence of target orbit altitude, payload mass, typical missions, and the type of liquid propellant (oxidizer - fuel pair) used on the launcher characteristics is quantified. The launchers are obtained using a multidisciplinary design optimisation (MDO) algorithm, where the lowest mass configuration capable of accurately inserting the payload into the target orbit is considered to be optimal.

In the Frame of Romanian Incentive Scheme Programme under European Space Agency, a feasibility study of a Small Orbital Launcher was performed by "Politehnica" University of Bucharest. Building on the results obtained in the project, European Space Agency's Space Transportation Directorate through the Future Launchers Preparatory Programme awarded INCAS a Phase 0/A contract for a Microlauncher concept. The targeted market is the micro-satellites sector developed by universities, but also by small companies for research purpose and validation of applications/products before commercialization, with a mass lower than 150 kg. This paper presents a conceptual design for a Microlauncher having a target mission to deliver a small satellite of 150 kg into a 600 km circular polar orbit with the launching site located in Europe. The Microlauncher concept is obtained using an in-house multidisciplinary optimization tool, which contains four modules: weights and sizing, aerodynamics performance assessment, propulsion, trajectory computation and optimization.