1,720,953 research outputs found
SAR Retracking in the Arctic: Development of a year round Arctic SAR retracker system
No full textThe sea level determination in the Arctic brings challenges with it, due to its high seasonality, varying surface types and scarce in-situ observations. By developing a year round retracker system with an implementation of a bias removal strategy, the Arctic sea level determination is improved. Through the analysis of the seasonal performance of retrackers, an insight into the behaviour of empirical and physical retrackers is gained that can be used in the selection process of optimal retrackers for given conditions and regions.SpaceflightAstrodynamics & Space MissionsAerospace Engineerin
Europan tidal deformation: Providing a theoretical framework for altimetry data to determine ocean presence
No full textAerospace Engineerin
The Unified State Model. Derivation and Applications in Astrodynamics and Navigation.
No full textAstrodynamics and Space MissionsAerospace Engineerin
The Mars Shuttle: An investigation into the feasibility of a shuttle vehicle between the Martian surface and an orbital node to support the continued presence of humans on Mars
No full textThe last 20 years has brought with it a surge in efforts toward the Red Planet as the next frontier is human space exploration draws closer. Many concepts have been proposed for a sustained human settlement on Mars, with NASA’s ISRU-to-the-wall campaign identifying the need for a shuttle vehicle between theMartian surface and a station in an orbit around the planet. Two such vehicle concepts have been designed: the Charon by Gaffarel et. al. and Hercules by Komar et. al.. However a Multidisciplinary Design Optimisation (MDO) has thus far not been applied. In this research an MDO is employed for the same mission scenario as the Charon vehicle’s. The vehicle must transport 1200 kg, including 6 crew, to a 607 km circular orbit at 44.96± inclination from the Martian base that is located at 42.5± North and 25.5± East. It must then return to the base, its entry beginning at 80 km altitude at a velocity of 3500 m/s. The MDO in this research is performed by dividing the design of the vehicle and its trajectories into various disciplines, which are optimised in parallel. In reality cost commonly is the dominant factor that drives the design, in this case, the vehicle’s Gross Take OffWeight (GTOW) is taken as the objective. Estimating the cost of the Martian shuttle vehicle within reasonable accuracy is exceedingly difficult, as not only is its realisation still decades away, costs such as the shuttle’s transportation to Mars, its operational costs, and its maintenance costs are very hard to estimate. However, a vehicle’s GTOW is directly influential on its cost, and its reduction is therefore the optimisation’s goal. The disciplines within the optimisation are the vehicle’s geometry, mass, aerodynamics, and ascent and descent trajectories. Constraints are set for both the vehicle and its trajectories, to which the design must adhere, and the performance of the design is determined by a fitness function that ensures the reduction of the vehicle’s GTOW. The optimiser itself is written using Tudat software, a set of C++ libraries developed by the TU Delft. As the same mission scenario is taken for the vehicle in this research as that of the Charon vehicle, the Charon design can be directly compared to that of the optimiser. TheMDO is able to significantly reduce the vehicle GTOW, obtaining an optimum solution of 146.2 tonnes, which is more than 20 tonnes lighter than Charon, at 168.1 tonnes. The greatest reduction in mass is found in the ascent propellant mass, which is the greatest contributor to the GTOW. This reduction is mostly due to the lower maximum Thrust to Weight (TW) ratio used in the MDO. Other scenarios are also investigated and their effects observed. Two other target orbits and rendezvous strategies are tested, namely the same as the Hercules vehicle (108 km pericentre altitude and an eccentricity of 0.0178), and a direct ascent to the orbital node at a circular 500 km altitude orbit. The Hercules vehicle scenario proved to be by far the most GTOW-preferable, with a GTOW of only 103.0 tonnes. The GTOW of the MDO solution found for the Hercules scenario is also less than that of the Hercules vehicle design, which is 162.8 tonnes, however the Hercules vehicle transports a payload mass of 5750 kg as opposed to the MDO’s 1200 kg, therefore the mass reductions cannot be solely attributed to the optimiser performance. It is clear from the results that the altitude of the initial target orbit is the greatest factor contributing to the GTOW, with a reduction in both payload mass andMartian base latitude also shown to reduce the GTOW. Neither an increase in the maximum acceleration constraint, nor a change in ascent thrust profile, were shown to have any benefit on the GTOW. The sensitivity of the optimum design with respect to uncertainties is assessed. The final pericentre and final eccentricity are both most sensitive to the final pitch node value, especially when taking interference with other variables into account. The inclination, however, was almost solely influenced by the first pitch angle value. The latitude and longitude are also almost solely influenced by a single variable, namely the flight path angle, and the final velocity and final pitch angle are both highly volatile to all variables when interference effects are taken into account. TheMDO model as a whole is found to be sensitive to changes in both aerodynamics coefficients and propulsive efficiency; an increase in aerodynamic coefficients adversely affects the GTOW, and an increase in propulsive efficiency benefits the GTOW, and vice versa.Aerospace Engineerin
A feasibility study on the recovery of Electron's first stage using a vertical landing approach
No full textThe small satellite market is rapidly growing and becomes more competitive by the day. In order to reduce launch costs, the concept of reusable launch vehicles wins popularity. Only few companies have managed to successfully recover and re-use their launch vehicle, of which SpaceX is the only one using a powered descent technique. This study focuses on applying that same technique on a small satellite launcher: Rocket Labs Electron. A 3-DOF trajectory optimization is performed to find the fuel optimal ascent and descent trajectories of the Electron. The original design of the Electron is adhered to as much as possible. The optimization is performed for a variety of Single- and Multi objective optimizers and different settings. Although this is only a feasibility study in which many assumptions are made, the results look promising for further research to be done.Aerospace Engineerin
Necessity of non-rigid body models for launch vehicles
No full textIn previous research, the necessity of non-rigid body simulators was stated to be a requirement for long, slender bodies. However, it remains unclear for which vehicle design parameters non-rigid body effects become destabilising. A set of 2+N linearly independent equations of motions can be combined in a single matrix to determine the motion of a flexible launch vehicle with N slosh masses. The flexibility of the vehicle is modelled using the method of assumed modes. The sloshing motion is modelled by assuming that the slosh mass behaves like a pendulum. To couple the different motions, a constraint matrix is used. The model is developed in Matlab Simulink. Using non-quiescent starting conditions with no external forces acting on the vehicle except gravity, it was found that the energy of the system remained constant. The flight data from the Stratos III launch vehicle was obtained and compared to the simulated non-rigid body model data. It was found that slosh motion can result in destabilisation when this vehicle experiences a sudden decrease in acceleration. The destabilisation that occurred could not be observed when rigid body equations of motion were used. Based on a sensitivity analysis over a range of vehicle parameters, it is recommended for length over diameter (L/D) ratios above 20 to consider the use of non-rigid body models. From this ratio, the maximum and total flexibility strongly increase. For L/D ratios below 20, flexibility is negligible and sloshing becomes the main perturbing factor. It was found that a PID controller that is designed using a rigid body model is still able to stabilise the non-rigid body during flight. Oscillations occurred in both the rigid and non-rigid body model when a single set of gains was used. For both models, the oscillations were removed by introducing gain scheduling.Aerospace Engineerin
Going Beyond Counting First Authors in Author Co-citation Analysis
Get PDFThe present study examines one of the fundamental aspects of author co-citation analysis (ACA) - the way co-citation
counts are defined. Co-citation counting provides the data on which all subsequent statistical analyses and mappings
are based, and we compare ACA results based on two different types of co-citation counting - the traditional type that
only counts the first one among a cited work's authors on the one hand and a non-traditional type that takes into
account the first 5 authors of a cited work on the other hand. Results indicate that the picture produced through this non-traditional author co-citation counting contains more coherent author groups and is therefore considerably clearer. However, this picture represents fewer specialties in the research field being studied than that produced through the traditional first-author co-citation counting when the same number of top-ranked authors is selected and analyzed. Reasons for these effects are discussed
Re-entry vehicle aerodynamic database reconstruction from the analysis of test dynamics
No full textDuring the design and analysis of a mission characterised by re-entry flight, consistent effort and resources are invested in the definition of the aerodynamic database of the re-entry vehicle. An aerodynamic database collects a set of dimensionless coefficients that describe the interaction of the vehicle's geometry and airflow, and an accurate estimate of its elements is key to the correct modelling of the accelerations experienced during flight.The methodology elaborated herein focuses on the reconstruction of the aerodynamic database of a capsule-shaped vehicle, based on the analysis of simulated data from wind tunnel tests and drop tests. The estimation process is characterised by linear regression of the data that requires linearisation of the dynamics, with the use of Taylor series expansions, and a polynomial representation of the coefficients' dependency on the angle of attack. The preliminary estimate of the coefficients computed by linear regressions is introduced in data smoothing models in order to reduce the noise and errors present in the dataset. The effectiveness of the Extended Kalman filter, the Unscented Kalman filter and the Square Root Unscented Information filter applied to the measurement is established and the results proved their performance to be comparable one to another in the presented problem.Finally, the core of the research performed is related to nonlinear regression methods, oriented towards aerodynamic database reconstruction. At first, an analytical approach is elaborated by defining a harmonic solution for curve fitting of oscillatory dynamics. However, the iterative Gauss-Newton algorithm does not converge to a definition of the regression parameters in the presented case. The focus is therefore then concentrated on optimisation methods. Multi-island Genetic Algorithm, Adaptive Simulated Annealing, Nealder & Mead Downhill Simple, Hooke-Jeves Direct Search and a Hybrid Algorithm are all methods applied to the problem. From the analysis of the results, based on a simulated wind tunnel test of a Hayabusa type capsule in subsonic flow regime, three algorithms emerge for greater accuracy in the reconstructed database: the Hybrid Algorithm, the Hooke-Jeves Direct Search and the Adaptive Simulated Annealing method.Aerospace Engineerin
Dawn Aerospace Mk-III: An exploration of cost driven mission scenarios of a winged Two Stage to Orbit semi-Reusable Launch Vehicle integrated in the common airspace
No full textThe demand of small payload launch vehicles has been growing over the past years. Customers base their selection of launch vehicles on cost-effectiveness, flexibility, availability and reliability. A new launch vehicle, the Dawn Aerospace Mk-III, is proposed to be developed, while designed to take into account all these criteria.For flexible and frequently available operations the launch trajectory is integrated in the common airspace. Manoeuvrability is identified as a key capability for safe operations in the common airspace. For high manoeuvrability the first stage is designed as a rocket propelled airplane, a so-called 'spaceplane'. The expendable upper stage is stored internally. After payload injection the first stage returns to the spaceport of take-off. This means the System is a Two Stage to Orbit semi-Reusable Launch Vehicle integrated in the common airspace.Cost-effectiveness is a primary selection criterium in the decision making of customers, which is why cost is included from an early stage in the development. This study shows that for identifying the cost gradient in the design space the total dry mass of the vehicle is sufficient. In this way cost optimality is included, although the Cost per Flight cannot be determined. What is determined is the effect of different technical and operational considerations. Taking into account qualitative cost differences, a selection of cost derived mission scenarios are studied. This includes different Return to Launch Site methods, first stage engine design and lay-out, the prohibition of fairing usage and integrated landing gear for take-off and landing.To analyse and optimize the different designs a Multidisciplinary Design Optimization tool is developed. This tool optimizes the vehicle and the ascent trajectory simultaneously to determine the lowest total dry mass solution meeting all requirements and constraints. To estimate the aerodynamic performance of the first stage the X-34 Advanced Technology Demonstrator is used as a reference vehicle. This means the geometry of the first stage is not optimized, while the size is. The upper stage is modelled as a conventional upper stage, of which the size and geometry is optimized.The study shows that the proposed design is feasible, meeting all requirements and constraints. The result is a vehicle with a total dry mass of 6273.0 kg, Gross Take-Off Weight of 42972.8 kg and a first stage length of 19.4 m. Of the total dry mass 94.5% is reusable. The return of the first stage is driving the trajectory design, as a steep ascent trajectory is required for limiting the downrange of the first stage. This results in 30%-50% more gravity loss in the System, which demands for a high ΔV performance. Due to the size of the first stage and the propellant required for returning, the first stage ascent ΔV is limited. For that reason, the upper stage design has a Propellant Mass Fraction of 0.939, increasing the upper stage ascent ΔV performance. Achieving such a Propellant Mass Fraction is possible, but challenging. The upper stage design is identified as a key element in the System performance for successfully meeting all mission requirements.Three different Return to Launch Site methods are compared. Two methods, in-plane pitch over and aeroturn, are active which requires return propellant. The third method, glideback, is passive. The study shows that glideback can be favourable. The total dry mass increase is 4.4% when compared to in-plane pitch over. The increase in heat load is ~16%, but the heat load in this study is limited with a total heat load of ~2.0 MJ/m^2. However, for glideback an even higher upper stage Propellant Mass Fraction of 0.946 is required. This means that the result of the upper stage design determines if this return method is feasible.Using a shared engine design on the first and upper stage shows promising results. Using a single first stage engine reduces the first stage dry mass by 4.0%. However, using a shared engine design is expected to decrease the development cost drastically, due to the reduced size of individual engines. The prohibition of fairing usage shows a stronger effect on the result. Allowing the use of a fairing decreases the first stage dry mass by a maximum of 9.6%. Fairing usage on the other hand harms the safe operations in the common airspace as the ejection of uncontrolled material requires large safety zones. The penalty on mass is acceptable for allowing integration in the common airspace. A landing gear sized for take-off results in a heavier vehicle. When the first stage is supported by a cart during take-off the first stage dry mass is decreased by 24.6%. When the first stage is air-launched the first stage dry mass decreases by 33.3%.Aerospace Engineerin
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