The calibration of humidity sensors is usually carried out in air and at near ambient pressure. In many industries, humidity sensors are sometimes used to make measurements in carrier gases other than air and this can give rise to errors. It is presumed that humidity sensors have responses that differ depending upon gas species–i.e. they are gas species dependent. The number of atoms in a gas molecule is called its Atomicity. A gas with a high atomicity can remove more heat out of a sensor than a gas with a small atomicity. The gases chosen in this research were methane and helium. The Flow Mixing Generator (FMG) has been developed to generate humid gas. The operation relies on the basis of mixing wet gas and dry gas. The major component was dry gas, >95% by volume. The experimental dew point temperature range was around -30 °C at a pressure of 1.5 bar A with this wet gas and dry gas ratio. The FMG was controlled to generate humid gases with the same dew point, although the dry gas‟ dew points were not equal. Therefore, the gas flow rates were adjusted. The mass flow controller that was a part of the FMG used to measure and control the gas flow rate was calibrated in air. When the other gases were used, the gas correction factor had to be applied in order to obtain the actual flow rate. The correction factor‟s uncertainty was not reported and presumably, this quantity may be included with the instrument‟s uncertainty (0.25% of value). Humidity quantities may be stated in various units. Water vapour is directly sensed by a humidity sensor. This quantity is thus important and usually reported in terms of water vapour pressure. Based on Dalton‟s partial pressure, total pressure is the sum of the pressures of gases in the mixture including water vapour pressure. However, this is only valid for gas at low pressure with no molecular interaction. The ideal condition does not exist – even in ambient air. The actual water vapour pressures need to be corrected by the enhancement factor. The present equations for the enhancement factor are valid for air at pressures up to 20 bar. The equations of states with respect to molecular interaction were chosen to calculate the enhancement factor. Nevertheless, the equations of state were insufficient. The overall uncertainty evaluation for the standard hygrometers, the FMG and the test humidity sensors have been presented. The FMG gives an acceptable uncertainty of ± 0.3 °C at the 95% confidence level (k = 2). Weak reproducibility gave the largest source of uncertainty. The polymer sensor showed a difference of 1.2 °C between air and methane, but the difference was insignificant in the aluminium oxide sensors. It can be concluded that gas species has an effect on a polymer sensor.

Recently proposed space missions such as Darwin, LISA and NGGM have encouraged the development of electric propulsion thrusters capable of operating in the micro-Newton (N) thrust range. To meet these requirements, radio frequency (RF) gridded ion thrusters need to be scaled down to a few centimetres in size. Due to the small size of these thrusters, it is important to accurately determine the thermal and performance parameters. To achieve this, an RF ion thruster model has been developed, composed of plasma discharge, 2D axisymmetric ion extraction, 3D electromagnetic, 3D thermal and RF circuit models. The plasma discharge model itself is represented using 0D global, 2D axisymmetric and 3D molecular neutral gas, and Boltzmann electron transport sub-models. This is the rst time such a holistic/comprehensive model has been created. The model was successfully validated against experimental data from the RIT 3.5 thruster, developed for the NGGM mission. Afterwards, the computational model was used to design an RF gridded ion thruster for an Ion Beam Shepherd (IBS) type space debris removal mission. Normally, the IBS method requires two thrusters: one for impulse transfer (IT) and one for impulse compensation (IC). This thesis proposes a novel thruster concept for the IBS type missions where a single Double-Sided Thruster (DST) simultaneously producing ion beams for the IT and IC purposes is used. The advantage of DST design is that it requires approximately half the RF power compared with two single-ended thrusters and it has a much simpler sub-system architecture, lower cost, and lower total mass. Such a DST thruster was designed, built and tested, with the requirements and constraints taken from the LEOSWEEP space debris removal mission. During the experimental campaign, a successful extraction of two ion beams was achieved. The thesis has shown that it is possible to control the thrust magnitudes from the IT and IC sides by varying the number of apertures in each ion optics system, proving that the DST concept is a viable alternative for the LOESWEEP mission.

ICMEs observed by the MESSENGER and Venus Express spacecraft have been catalogued and analysed. The ICMEs were identified by smooth rotations of the B-field direction consistent with a flux rope structure, coinciding with an enhanced field strength. Thirty-five ICMEs were found in the surveyed MESSENGER data (~March 2007 to April 2012), and eighty-four ICMEs in the surveyed Venus Express data (May 2006 to December 2013). Ropes with northward leading edges were four times more common than ropes with southward leading edges, in agreement with a previously established solar cycle dependence. Ropes at low inclinations to the solar equatorial plane were four times more common than high-inclination ropes, possibly an observational effect. In addition, data from MESSENGER, Venus Express, STEREO and ACE were examined for multi-point signatures of the catalogued ICMEs. For spacecraft separations < 15° in heliocentric longitude, the second spacecraft observed the ICME’s rope in 82% of cases; this percentage fell to 49% for 15–30° separations, to 18% for 30–45° separations, and to 12% for 45–60° separations. One ICME, observed by MESSENGER and STEREO-B while the spacecraft were radially aligned, has been analysed in detail. Few such radial observations of ICMEs have been reported previously. Force-free fitting indicates that the ICME’s rope diameter increased with an r_H^0.94 dependence on heliocentric distance, and the axial B-field strength dropped with an r_H^-1.84 dependence, clear indications of expansion. Axial magnetic flux was conserved, suggesting that the rope underwent no significant erosion through reconnection. The change in the rope’s angular width has been estimated via helicity conservation. The rope axis rotated by ~30° between the spacecraft towards the solar equatorial plane. A preliminary study of a larger number of radially-observed ICMEs indicates that rope rotations of this magnitude are not uncommon; such rotations would have important implications for space weather forecasting.

This thesis gives the conceptualization of inflation of inflatable membrane space structures. Although there has been little study using software simulation and the majority of documented research is based on theoretical numerical calculations. This research advanced the prior understanding of wrinkling within inflated membranes by using complex structures subjected to deformation loads. Within this thesis, a computational framework for the numerical analysis of the interaction between acting forces on the membrane and the membrane structure dynamics is presented. Moreover, in the case with thin membrane deformations, the synergy between the membrane wrinkling and structural forces has to be examined. This membrane structure-anatomical forces correlation results in a dynamic wrinkling problem, which can only be modelled easily and effectively by a simulation software that can integrate each assumption and attribute within the analysis. In the structural simulation within Abaqus FEA software, key consideration has to be given in modelling the geometric non-linearity behaviour of the membrane. By using the existing continuum expression for the virtual internal work in curvilinear coordinates. This is used to derive the modified fundamental formulation in which all subsequent analysis is established on and the initial equilibrium shape of the membrane. A critical feature of the new formulation is the prospect of adding pre-stressed forces to the membrane structure. The approach developed, established on an alteration of the material stiffness matrix to integrate the effects of wrinkling and deformation, can be utilized to calculate the behaviour of the membrane within a finite element simulation. In the wrinkling model, the state of the membrane element (taut, wrinkled or slack) is characterized by a mixed wrinkling criterion. Once it has been identified that the membrane element is wrinkled, an iterative scheme looks for the wrinkled orientation angle and the precise stress distribution, including only uni-axial tension in the wrinkle direction, is then derived. The wrinkling model has been verified and validated by contrasting the simulated conclusions with documented results for the instance of a time-independent isotropic membrane subjected to shear and axial loading. Utilizing the time integration method, a time-dependant pseudo-elastic stiffness matrix was represented and therefore, rather than calculating the convolution integral all through the Abaqus simulation, then we can calculate the behaviour of a membrane structure by superposition of a series of step by step increments in basic finite element software. The theoretical computations from the Abaqus/Explicit analysis were compared with documented results for the shear and axial loading. The results agreed very well, assuming friction and any relativistic dynamic effects were excluded. The discrepancy between the shear loading solution is 7% while the discrepancy between the axial loading is only 5% between the Abaqus modeland the documented model. This discrepancy could be the resultant of the source of energy dissipation from the visco-elastic behaviour during the loading and unloading of forces. It can be stated that for the Kapton HN membrane, this result falls within acceptable range but to increase accuracy, the load and unloading will be carried out on a set steady amplitude to inhibit in shock effects within the model. A three-dimensional finite element model which integrates wrinkling and frictionless contact has been developed to simulate the adaptive smart cell and cylindrical membrane structure. The loading of both structures is given by a non-uniform differential inflation pressure with a continual gradient adjacent to height. The resultant solutions are computed using Abaqus/Explicit software, with an integrated user-defined material subroutine to account for elastic wrinkling deformation that administers a combined stress-strain criterion. Frictionless contact within the membrane structure is prescribed for both complex structures (Adaptive Smart Structures Model and Inflatable Beam Model) in order to prohibit the penetration of the membrane structure through itself. Both the complex inflatable membrane wrinkling models accomplish the purpose of exceptional subgrid scale performance in relation to accuracy, competency, computing hardware & software expense, complexity and the model convergence rate. The numerical algorithm is created in general context and is flexible for a large variety of material models. For a closed membrane structure, the skew symmetric constraint parameters vanish, while the existing symmetric domain variables mirror preservation of the system. This procedure does not demand the discretization of the fluid (gas) domain or the link between coupling of fluid (gas) and membrane. As a result of this basic fact, the computation is drastically simplified. The adaptive structures model introduces a novel approach in harnessing solar power for reuse on the ground as a stable source of power. The simulations were based on the space part of the stiff structure created of hexagonal membrane cells. Simulations are carried out in Abaqus Finite Element Analysis software for simplicity & a comparison for validation purposes is tested against an experimental inflatable cell within a vacuum chamber. It was showcased that the final configuration could be achieved regardless of the packaging shape of the inflatable cell array. The inflatable beam model is comprised of two sections, the bending & buckling of the inflated beam and the post-inflation of the bent and buckled beam. Abaqus software was used to simulate the inflatable beam during each configuration utilizing the integration of a modified VUMAT subroutine. A comparison is showcased representing the importance of the integration of the VUMAT subroutine within our Abaqus model.

No description available.

No description available.

Nowadays, space structures are often designed to serve only a single objective during their mission life, examples range from solar sail for propulsion over shields for protection to antennas and reflectors for communication and observation. By enabling a structure to deploy and change its shape to adapt to different mission stages, the flexibility of the spacecraft can be greatly increased while significantly decreasing the mass and the volume of the system. Inspiration was taken from nature. Various plants have the ability to follow the sun with their flowers or leaves during the course of a day via a mechanism known as heliotropism. This mechanism is characterized by the introduction of pressure gradients between neighboring motor cells in the plant’s stem,enabling the stem to bend. By adapting this bio-inspired mechanism to mechanical systems, a new class of smart deployable structures can be created. The shape change of the full structure can be significant by adding up these local changes induced by the reoccurring cell elements. The structure developed as part of this thesis consists of an array of interconnected cells which are each able to alter their volume due to internal pressure change. By coordinated cell actuation in a specific pattern, the global structure can be deformed to obtain a desired shape. A multibody code was developed which constantly solves the equation of motion with inputs from internal actuation and external perturbation forces. During the inflation and actuation of the structure, the entities of the mass matrix and the stiffness matrix are changed due to changing properties of the cells within the array based on their state and displacement. This thesis will also give an overview of the system architecture for different missions and shows the feasibility and shape changing capabilities of the proposed design with multibody dynamic simulations. Furthermore, technology demonstrator experiments on stratospheric balloons and sounding rockets have been carried out to show the applicability and functionality of the developed concepts.

Surrey Space Centre (SSC) has been working on an autonomous fixed-wing all-electric vertical take-off and landing (VTOL) aerobot for the exploration of Mars for several years. SSC’s previous designs have incorporated separate vertical lift and horizontal pusher rotors as well as a mono tilt-rotor configuration. The Martian aerobot’s novel Y-4 tilt-rotor (Y4TR) design is a combination of two previous SSC designs and a step forward for planetary aerobots. The aerobot will fly as a Y4 multi-rotor during vertical flight and as a conventional flying wing during horizontal flight. The more robust Y4TR configuration utilizes two large fixed coaxial counter rotating rotors and two small tilt-rotors for vertical takeoff. The front tilt-rotors rotate during transition flight into the main horizontal flight configuration. The aerobot is a blended wing design with the wings using the "Zagi 10" airfoil blended to a center cover for the coaxial rotors. The open source design and analysis programs XROTOR, CROTOR, Q-BLADE, XFLR5, and OpenVSP were used to design and model the aerobot’s four rotors and body. The baseline mission of the Y4TR remains the same as previously reported and will investigate the Isidis Planitia region on Mars over a month long period using optical sensors during flight and a surface science package when landed. During flight operations the aerobot will take off vertically, transition to horizontal flight, fly for around an hour, transition back to vertical flight, and land vertically. The flight missions will take place close to local noon to maximize power production via solar cells during flight. A nonlinear six degree of freedom (6DoF) dynamic model incorporating aerodynamic models of the aerobot’s body and rotors has been developed to model the vertical, transition, and horizontal phases of flight. A nonlinear State-Dependent Riccati Equation (SDRE) controller has been developed for each of these flight phases. The nonlinear dynamic model was transformed into a pseudo-linear form based on the states and implemented in the SDRE controller. During transition flight the aerobot is over actuated and the weighted least squares (WLS) method is used for allocation of control effectors. Simulations of the aerobot flying in different configurations were performed to verify the performance of the SDRE controllers, including hover, transition, horizontal flight, altitude changes, and landing scenarios. Results from the simulations show the SDRE controller is a viable option for controlling the novel Y4TR Martian Aerobot.

One means to reduce both the cost and the risk associated with space missions is to employ a vehicle designed within the re-usable, airliner-like 'space plane' paradigm. Key to the practicality of such vehicles is the further development of Combined Cycle Propulsion technology. In this thesis, a new engineering tool called the HYbrid PRopulsion Optimizer (HyPro) is presented, with the aim of analysing the performance of diverse engine concepts. The tool is conceived to be modular and flexible, and makes use of parsimonious modelling, in order to describe the engine at an high level of abstraction and to be fast in execution. A configurational optimizer has also been developed in order to automatically generate new design concepts, optimizing the engine cycle structure. It is conceived to be used at the beginning of development in order to perform an automatic and objective trade-off of possible propulsion solutions. In this work the model has been implemented for Rocket-Based Combined Cycle, and it has been verified and validated against analytical models, computational fluid dyanamic analyses and experimental data. The design proposed by the optimizer in these conditions was coherent with manually designed Combined Cycle Propulsion engines, demonstrating the HyPro's capability to converge on good solutions. The results, although preliminary, are very promising and represent a novelty in the field, since a configurational optimization, in the field of propulsion concepts, has never been attempted before. The results presented here demonstrate that the configurational optimization of engine design is viable. The next steps to produce a practical optimizer, which delivers robust and innovative engine solutions, are the addition of modelling capabilities beyond the Rocket-Based Combined Cycle and analysis discipline beyond the pure performances.

The requirement for satellites to have a mitigation or deorbiting strategy has been brought about by the ever increasing amount of debris in Earth orbit. Studies have been used to formulate space debris mitigation guidelines, and adherence to these guidelines would theoretically lead to a sustainable environment for future satellite launches and operations. Deployable sail designs that have traditionally been studied and used for solar sails are increasingly being considered for de-orbit applications. Such sail designs benefit from a low mass and large surface area to achieve efficient thrust. A sail has the potential to be used for drag augmentation, to reduce the time until re-entry, or as an actual solar sail – to deorbit from higher orbits. A number of concerns for sail-based deorbiting are addressed in this thesis. One of these concerns is the ability of a sail to mitigate the risk of a collision. By investigating both the area-time-product (ATP) and collision probability it is shown that a gossamer sail used for deorbiting will lead to a reduction in overall collision risk. The extent to which the risk is reduced is investigated and the contributing factors assessed. Another concern is that of attitude stability of a host satellite and deorbit sail. One of the biggest benefits of drag augmentation is the fact that it can achieve the deorbiting goal with an inactive host satellite. There is thus no need for active control, communications or power after deployment. But a simple 2D sail will lose efficiency as a deorbiting device if it is not optimally oriented. It was found in this research that it is possible for a host satellite with attached sail to maintain a stable attitude under passive conditions in a drag deorbiting mode. Finally, in order to fully prove the benefit of sail-based deorbiting it is shown that in certain scenarios this alternative might be more efficient at reducing collision risk, weighs less, and has less operational requirements than other alternatives such as electrodynamic tethers and conventional propulsion. This thesis aims to cover the fundamental concerns of a sail-based deorbiting device at mission level by firstly addressing the mission analysis aspects and then applying it to specific scenarios. The theory and methods required to perform mission analysis for a sail-based deorbiting strategy is presented. These methods are then used to demonstrate passive attitude stability for a drag sail, and reduction in collision risk, both in terms of the Area-Time-Product and collision probability. The analysis results are then further applied by identifying scenarios to which the proposed deorbiting device applies, and then performing a meaningful comparison by analysing a number of case studies. The application is made more concrete by comparison with likely contenders – traditional propulsion, electrodynamic tethers and an inflatable sphere.