1,721,039 research outputs found
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
No full textImpact 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
Analysis of the plasma transport in numerical simulations of helicon plasma thrusters
Full text linkThe accurate simulation of the plasma transport in helicon sources is a key aspect to improve the design of Helicon Plasma Thrusters (HPTs). Specifically, the 3D-VIRTUS code was proven to provide satisfactory estimations of the propulsive performance of realistic HPTs (difference between measures and numerical estimations of the thrust <30%). Nonetheless, further investigations are needed to deepen the influence that the plasma chemistry model, the formulation of the energy equation, and the definition of the diffusion coefficients have on the results of the simulation. First, a quantitative analysis has been conducted on a simplified configuration of HPT to study each phenomenon separately. Second, the generalized fluid model has been benchmarked against measures of plasma density performed on a helicon source. The radiative decay reactions affect the estimation of the performance (e.g., thrust) up to 40%. The quasi-isotherm formulation of the energy equation affects results (e.g., electron density) up to 30%. Accounting for anomalous transport or defining diffusion coefficients classically does not have a major effect on the simulation (e.g., thrust varies less than 20%). The generalized formulation of the fluid model provides estimations of the plasma density, which are within the uncertainty band of the measures (i.e., differences <20%)
Development of a lumping methodology for the analysis of the excited states in plasma discharges operated with argon, neon, krypton, and xenon
Full text linkIn this paper, a methodology is presented to compute the plasma properties (e.g.,, density and temperature) accounting for the dynamics of the excited states. The proposed strategy applies to both zero-dimensional (0D) models and multidimensional fluid and hybrid codes handling low-pressure (<50 mTorr) plasma discharges filled with argon, neon, krypton, and xenon gases. The paper focuses on two main aspects: (i) a lumping methodology is proposed to reduce the number of reactions and species considered in order to keep at bay the computational cost without a major loss of accuracy; (ii) the influence that different datasets of cross sections have on the results has been assessed. First, the lumping methodology has been implemented in a 0D model accounting for singly charged ions, neutrals, along with 1s and 2p excited states (Paschen notation). Metastable and resonant are treated as two separate species within the 1s energy level respectively). The results have been benchmarked against those obtained treating each energy level of the excited states as an individual species. Differences lower than 1% have been obtained. Second, the results of the 0D model have been compared against measurements of electron density and temperature performed on an inductively coupled plasma. Numerical predictions and experiments present a disagreement up to 20%-30%, which is comparable to the uncertainty band of the measurements. Finally, the lumping strategy has been implemented in a 2D fluid code to assess its computational affordability, and the results have been compared against the experiments as well. A variance up to 30% in electron density and temperature is registered adopting different datasets of cross sections
Investigation of different strategies for access to space of small satellites on a defined LEO orbit
Full text linkIn recent years small satellites have shifted from being secondary items to dominate the space market thanks to, but not only, the development of the large LEO constellations. For small satellites to be highly effective, particularly when arranged in such constellations, each satellite has to be placed in a specific orbital plane and orbital position. Currently, three main options are available to reach a specific orbit. The first, more straightforward solution is to employ a dedicated launch with a small launch vehicle. Plenty of small launchers are in development around the world (with few already operational); however, their cost per kilogram is predicted to be much higher than for large launchers. The second possibility is the exploitation of a rideshare option, where the satellite is transported by a large launcher together with other payloads on a general predetermined orbit. Afterwards the satellite needs to be transferred to its designed orbit. This can be done in two ways: through the use of a satellite carrier (also called self-propelled dispenser) or with the satellite own propulsion system. In both last cases, the mass transported by the launcher into the release orbit is higher than the final one necessary for the nominal mission, impacting total costs. In this paper these three possibilities are compared considering the need to reach various specific orbits, starting from a different release one in the case of the rideshare options. First of all, the change in velocity for different orbital parameters (altitude, eccentricity, phase, argument of perigee, inclination, RAAN) is computed. Afterwards the propulsion mass budget is calculated. Everything else being equal, it is demonstrated mathematically that a dispenser is inherently less efficient than a group of autonomous satellites, particularly for the RAAN change and a large number of carried satellites. However, it is not always possible or convenient to provide small satellites, particularly the smallest ones (nanosats/cubesats/microsats) with comparable propulsion capabilities of a larger dispenser, making the latter still an attractive option in several situations. A cost analysis also shows that, particularly for sophisticated small satellites, when the final orbit is far from the release one, a dedicated small launch vehicle can be cost competitive with the nominally much cheaper large launcher
Numerical analyses of thermal protection design in hybrid rocket motors
No full textHybrid rockets can provide a similar degree of flexibility in the thrust profile as a liquid engine but keeping a much simpler architecture. In addition to deal with only a single liquid propellant, hybrids are often designed with some sort of ablative protection system borrowed from the less flexible solid rockets instead of the more complex regenerative-type typically found in liquid rockets. Even if it is not a strict rule as regenerative hybrids or ablative liquids exist, this is the most common approach. In order to properly design the hybrid combustion chamber, it is important to determine the behavior of the thermal protection system during the burn and after the burn, particularly when the multiple fire capability of hybrids has to be exploited. In this paper, the thermal response of thermal protections is investigated trough numerical modelling. A one-dimensional code has been developed that solves the transient heat equation with or without regression of the surface. The code considers the heat transfer normal to the surface from the combustion trough the thermal protection up to the external surface of the hybrid casing, where the radiative heat transfer toward space is applied. The results highlight the importance of the heat soak back after burn, which force the use of thicker thermal protections, higher temperature resistant materials and more careful design if the hybrid is fired multiple times or when the motor case is foreseen to be reusable. However, it is also shown that, when possible, properly using the thrust termination and re-ignition capability of hybrids can help limiting the amount of thermal protections to a level even lower than that of the single burn expendable case. Nevertheless, on the opposite side, other critical situations like an upper stage performing an Hohmann transfer are also highlighted. The methodology and the analyses performed in this paper can also be applied/extended to non-regenerative cooled liquid engines
Investigation of different strategies for access to space of small satellites on a defined LEO orbit
No full textIn recent years small satellites have shifted from being secondary items to dominate the space market thanks to, but not only, the development of the large LEO constellations. For small satellites to be highly effective, particularly when arranged in such constellations, each satellite has to be placed in a specific orbital plane and orbital position. Currently, three main options are available to reach a specific orbit. The first, more straightforward solution is to employ a dedicated launch with a small launch vehicle. Plenty of small launchers are in development around the world (with few already operational); however, their cost per kilogram is predicted to be much higher than for large launchers. The second possibility is the exploitation of a rideshare option, where the satellite is transported by a large launcher together with other payloads on a general predetermined orbit. Afterwards the satellite needs to be transferred to its designed orbit. This can be done in two ways: through the use of a satellite carrier (also called self-propelled dispenser) or with the satellite own propulsion system. In both last cases, the mass transported by the launcher into the release orbit is higher than the final one necessary for the nominal mission, impacting total costs. In this paper these three possibilities are compared considering the need to reach various specific orbits, starting from a different release one in the case of the rideshare options. First of all, the change in velocity for different orbital parameters (altitude, eccentricity, phase, argument of perigee, inclination, RAAN) is computed. Afterwards the propulsion mass budget is calculated. Everything else being equal, it is demonstrated mathematically that a dispenser is inherently less efficient than a group of autonomous satellites, particularly for the RAAN change and a large number of carried satellites. However, it is not always possible or convenient to provide small satellites, particularly the smallest ones (nanosats/cubesats/microsats) with comparable propulsion capabilities of a larger dispenser, making the latter still an attractive option in several situations. A cost analysis also shows that, particularly for sophisticated small satellites, when the final orbit is far from the release one, a dedicated small launch vehicle can be cost competitive with the nominally much cheaper large launcher
Semi-Analytical Model of a Helicon Plasma Thruster
No full textThis article is devoted to the presentation of a semianalytical model of a helicon plasma thruster based on: 1) a 0-D global source model (GSM) for the simulation of the plasma source; 2) a monodimensional acceleration model for the plasma expansion and thrust production in the magnetic nozzle; and 3) a detachment criterion for the identification of the location in the nozzle where the plasma separates from the magnetic field lines. The developed GSM permits the simulation of magnetic topologies provided for radial cusps in the source region. Three different acceleration models presented in the literature have been implemented and employed for component 2) of the full thruster model. Two detachment criteria have also been considered for component 3). A comparison between the characteristics and the results of the different models employed has been carried out. Finally, the results obtained by all the combinations of acceleration models and detachment criteria presented have been compared against the solutions of a numerical fluid code and experimental thrust measurements. A maximum relative deviation between the results of the acceleration models of 3% was found, while a thrust variation of up to 20%-30% was recorded by changing the detachment criteria. A maximum relative error between the model's predictions and the experimental measurements of 25% was obtained
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