1,720,999 research outputs found
Review of Alternative Sustainable Fuels for Hybrid Rocket Propulsion
Full text linkHybrid rockets using specific oxidizer–fuel combinations are considered a green alternative to current propulsion systems, as they do not release very toxic or polluting exhausts, but only much less harmful substances such as carbon monoxide/dioxide and soot. However, in a long-term vision where space access and rocket transportation become a daily routine all around the world, the simple use of current green propellants could begin to become insufficient if the rest of the industry already follows much stricter rules, which are expected to tighten significantly in the future, thereby making emissions from rocket flights no more negligible. In this paper, the possible use of alternative sustainable solid fuels for hybrid rockets that are not derived from fossil fuels and are ideally carbon neutral is investigated and discussed based on the available data in the hybrid literature and on the literature related to renewable fuels in general. Even if this topic is apparently far away from the current necessities, as hybrid propulsion is not yet operational, it is paramount to consider a long-term vision and associated research efforts to make sure that the potential hybrid propulsion introduction to the commercial market is more than a simple flash in the pan, but offers a solid opportunity
Challenges of Ablatively Cooled Hybrid Rockets for Satellites or Upper Stages
Full text linkAblative-cooled hybrid rockets could potentially combine a similar versatility of a liquid propulsion system with a much simplified architecture. These characteristics make this kind of propulsion attractive, among others, for applications such as satellites and upper stages. In this paper, the use of hybrid rockets for those situations is reviewed. It is shown that, for a competitive implementation, several challenges need to be addressed, which are not the general ones often discussed in the hybrid literature. In particular, the optimal thrust to burning time ratio, which is often relatively low in liquid engines, has a deep impact on the grain geometry, that, in turn, must comply some constrains. The regression rate sometime needs to be tailored in order to avoid unreasonable grain shapes, with the consequence that the dimensional trends start to follow some sort of counter-intuitive behavior. The length to diameter ratio of the hybrid combustion chamber imposes some packaging issues in order to compact the whole propulsion system. Finally, the heat soak-back during long off phases between multiple burns could compromise the integrity of the case and of the solid fuel. Therefore, if the advantages of hybrid propulsion are to be exploited, the aspects mentioned in this paper shall be carefully considered and properly faced
Optimal use of electric propulsion for drag compensation in very low earth orbit on satellites with deployable solar panels
Full text linkIn recent years, thanks to advancements in commercial hardware technology miniaturization, small satellites have become more and more important and effective in the Space ecosystem. In previous papers it has been demonstrated that future smallest observation systems, operating at a lower altitude than traditional systems, have the potential for comparable or better performance, much lower overall mission cost, lower risk, shorter schedules, lower up-front development cost, more sustainable business model, mitigating the problem of orbital debris. Flying in very low Earth orbit requires the addition of a propulsion system capable of providing drag compensation. Thanks to its high specific impulse, electric propulsion is potentially well suited for this task. However, the use of an electric thruster often requires significant power that can be provided only with deployable solar panels. To minimize the drag, the solar panels should be aligned with the satellite orbital ve-locity. At the same time, to maximize the power input, the solar panel surfaces should point toward the Sun. In the majority of cases these two conditions are not met simultaneously, so an optimal trade-off is needed. This topic is investigated in this paper by the use of an orbital model of a generic satellite with deployable solar panels. The model is able to simulate any circular orbit of the satellite in any day of the year and calculate both the power input and the drag from the solar panels based on their inclination. An equivalent specific impulse is also defined taking into account the thrust that the electric thruster has to provide only to push the area of the solar panels required to power the propulsion system. The results show, as expected, that the sun-synchronous Dawn-Dusk (or Dusk-Dawn) orbit is the preferred one for the electric propulsion as the velocity vector and the Sun vector are almost orthogonal. In the other cases, at very low Earth orbit, even relatively small in-clinations of the solar panels respect to the velocity vector have a significant impact on the drag produced, increasing the propellant consumption and the required power. Nevertheless, it is shown also that if the solar panels are aligned with the velocity, the power lost can be kept reasonable. The same analysis could also apply to a satellite in low Earth orbit that does not employ an electric thruster but still needs deployable solar panels because of other sources of high power consumption (e.g. the payload)
Comparison between Different Re-Entry Technologies for Debris Mitigation in LEO
Full text linkThe population of satellites in Low Earth Orbit is predicted to growth exponentially in the next decade due to the proliferation of small-sat constellations. Consequently, the probability of collision is expected to increase dramatically, possibly leading to a potential Kessler syndrome situation. It is therefore necessary to strengthen all the technologies required for collision avoidance and end-of-life disposal of new satellites, together with active debris removal of current and potential future dead satellites. Both situations require the lowering of the altitude of a satellite up to re-entry. In this paper several de-orbiting technologies are evaluated: natural decay, chemical propulsion (solid and liquid), electric propulsion, drag sail, electrodynamic tether, and combinations of the previous ones. The comparison considers the initial altitude, system mass, de-orbiting time, collision probability during descent, reliability, and technological limits. Differences between active debris removal and satellite end-of-life self-disposal are taken into account. Moreover, the different types of re-entry, controlled vs. non-controlled, expendable vs. reusable system, demisable vs. non-demisable system are also discussed. Finally, the possibility to operate the satellite in Very Low Earth Orbits with a propulsion system for drag compensation and passive re-entry at end of life is investigated
Modeling and CFD Simulation of Regression Rate in Hybrid Rocket Motors
Full text linkAs the research on hybrid rocket motors advances, more accurate tools are needed to estimate the performance of the system by determining its fundamental parameters. One of them is certainly the regression rate of the solid fuel. Unfortunately, it depends on many complex physical phenomena and interactions which vary with time, space and scale, making the task of predicting its evolution very difficult. To address this issue, Computational Fluid Dynamics (CFD) was employed to investigate the inner workings of a hybrid rocket motor and develop a useful tool to help the design process and contribute to the physical understanding of the problem. By implementing a User-Defined Function (UDF) in a commercial CFD software, it has been possible to simulate the regression rate as a function of heat flux at the fuel surface. The calculation is performed by solving the energy balance at the solid-fluid interface coupled with the pyrolysis Arrhenius equation. Validation has been performed using literature data from Carmicino and Sorge. The results generally agree with the experimental regression rates within 10% of error for HDPE and 20% for HTPB. A significant discrepancy in the regression rates of these two fuels not accounted for by the classical theory was exposed
Preliminary Design of the Propulsion Subsystem for the European Advanced Reusable Satellite (EARS)
No full textThe European Advanced Reusable Satellite (EARS) is a European Commission funded program that aims at designing a reusable microsat to deliver commercial, scientific and institutional payloads to LEO and retrieving them back safely on Earth. The program seeks to develop the key enabling technologies for the foreseen mission. In this frame, T4i is designing the propulsion subsystem for the satellite and is developing the HTP and propane fed bipropellant thruster. The propulsion subsystem is responsible for delivering a precise de-orbit burn using 4 engines that rely on differential steering to keep the thrust vector aligned with the velocity vector. Differential steering is obtained through the thrusters’ pulsed mode where their nominal bi-propellant working mode is switched temporarily to a mono-propellant one by shutting off the fuel feeding valve. While this procedure slightly reduces the average specific impulse delivered by the platform, it allows for a simple and robust two axis control of the spacecraft during firing. A third controlling element, a propane cold gas reaction control system, was identified as the best solution for controlling rotations around the roll axis
Explicit analytical equations for single port hybrid rocket combustion chamber sizing
No full textHybrid rockets have several advantages with respect to current propulsion systems like simplicity, safety, reliability, environmental friendliness, and lower cost. To size the combustion chamber, it is fundamental to understand how the length, the external diameter, the volume loading, and the length-to-diameter ratio vary with the design parameters like scale, burning time, average mixture ratio, initial oxidizer flux, and propellant combination. The equations available in the literature are not in explicit form with respect to the aforementioned design parameters, and sometimes they can be misinterpreted by hybrid rocket engineers. Moreover, it is not possible to determine the instantaneous and average characteristic velocities during the burn without a numerical time integration. To show explicitly the real trends, a set of analytical equations has been developed. The key step is the definition of the relation between initial and average mixture ratio and the asymptotic treatment with respect to the ratio between the external and internal diameters. Moreover, an approximate explicit semi-analytical expression of the instantaneous and average characteristic velocities is provided. The explicit analytical equations are validated with the exact implicit solutions showing good agreement and exact asymptotic behavior
Investigation of the influence of gas phase molecular mass variations on hybrid rocket regression rate
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Impact of propulsion system characteristics on the potential for cost reduction of earth observation missions at very low altitudes
Full text linkEarth observation is one of the most important satellites’ applications. Past earth observation systems have used traditional space technology to achieve the best possible performance, but have been very expensive. Recently, thanks to advancements in technology and modern microelectronics, small satellites have become more and more useful at much lower costs, even if with reduced performance. The resolution of the optical payload improves as the altitude is reduced. Space system mass is proportional to the cube of the linear dimensions. This means that by flying at lower altitudes, satellites can reduce their payload size and therefore the entire mass of the satellite, thus reducing the cost of the system dramatically. However, almost all the earth observation missions fly at the minimum altitude that provides a sufficient orbital life. The addition of a propulsion system capable of providing drag compensation for the entire satellite operative life provides the possibility to fly at very low earth orbit. In this way, the same performance can be obtained with a smaller and cheaper system. To obtain the same coverage more units are needed to replace a larger unit at higher altitude. In this paper it is confirmed that future smallsat observation systems, operating at a lower altitude than traditional systems, have the potential for comparable or better performance, much lower overall mission cost (by a significant factor), lower risk (both implementation and operations), shorter schedules, lower up-front development cost, more sustainable business model, to be more flexible and resilient, more responsive to both new technologies and changing needs, and to mitigate the problem of orbital debris. This paper focus in particular on the effect of the propulsion system parameters (performance and costs) on the cost model as a function of the altitude. It is demonstrated that new affordable chemical propulsion systems provide already significant benefits with limited constraints, allowing a useful reduction of altitude and, consequently, costs. Electric propulsion systems have the potential to allow even lower altitudes or longer lifetimes; however, they have a stronger impact on the satellite design related to their power consumption, generally requiring deployable solar panels, which can limit the flexibility in the orbit selection or the added weight and cost of batteries. The development of electric thrusters that have good performance and limited impact on the satellite architecture (particularly at small scales) is fundamental to exploit their potential for reduced mission costs through very low altitude flight
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