Design of a reconfigurable satellite constellation

This paper provides a fully analytical method to describe a satellite constellation reconfiguration manoeuvre. By making use of low-thrust propulsion and exploiting the Earth’s natural perturbing forces it is possible to analytically describe the reconfiguration of a constellation, achieving a desired separation of both Right Ascension of Ascending Node (RAAN) and Argument of Latitude between satellites. An inherent trade-off exists between the time taken for a manoeuvre and the required ΔV, however the analytical solution presented here allows for a rapid visualisation of the trade-space and determination of the ideal transfer trajectory for a given mission. The general method presented can be applied across a range of scenarios, including constellation deployment and repurposing. The results show that for a scenario with an initial orbit semi-major axis of 6878.14km, and a desired final semi-major axis of 6778.14km it is possible to achieve a separation of 180° argument of latitude between a manoeuvring and a non-manoeuvring reference satellite in approximately 68 hours with a ΔV of 200m/s. To achieve the maximum possible RAAN separation of 90° with a ΔV of 200m/s requires a much longer time of over 218 days. Using two manoeuvring satellites with the same total manoeuvre ΔV was found to be more efficient only for short manoeuvre times. This is quantified and for the case considered it is found that using a 2-satellite manoeuvre is advantageous when changing the argument of latitude and when changing the RAAN <10° approximately. The ability to identify this turning point clearly is a distinct advantage of the analytical solution presented

General perturbation method for satellite constellation deployment using nodal precession

The dawn of "New Space" in recent years is changing the landscape of the space industry. In particular, the shift to smaller satellites, requiring shorter development time s and using off-the-shelf-components and standardized buses, has led to a continuing reduction in spacecraft cost. However, launch costs remain extremely high and frequently dominate the total mission cost. Additionally, many small satellites are designed to operate as part of a larger constellation, but traditional launch methods require a difference dedicated launch for each orbit plane to be populated. This need for multiple costly launches can stifle, and even prohibit, some missions requiring numerous orbit planes as the launch cost increases beyond what can be justified for the mission. As of 2014, most smallsats, including CubeSats, have been launched on opportunistic ‘rideshare’ or ‘piggy-back’ launches, in which the spacecraft shares its launch with other craft, often as a secondary payload. This has the advantage of providing a cheaper launch but restricts the operator’s choice of orbit, which will affect the system performance

An analytical low-cost deployment strategy for satellite constellations

This work proposes a novel method for the deployment of a constellation of nano-satellites into Low Earth Orbit by using carrier vehicles to deliver the nano-satellites into the required orbit positions. The analytical solution presented allows for rapid exploration of the design space and a direct optimisation of the deployment strategy to minimise the time for complete constellation deployment. Traditionally, the deployment of satellite constellations requires numerous launches – at least one per orbital plane – which can be costly. Launching as a secondary payload may offer significant cost reductions, but this comes at the price of decreased control over the launch schedule and final orbit parameters. The analytical method presented here allows for the optimal positioning of the orbit planes of the constellation to be determined and the minimum time for deployment determined as a function of the manoeuvre ΔV. The effect of atmospheric drag on the manoeuvre propellant cost is also considered to ensure a realistic deployment scenario. A case study considering three constellation designs is presented which compares the cost of deployment using traditional launch methods with that of deploying the constellation using carrier vehicles. The results of this study show a significant reduction in cost when using the carrier vehicles on a dedicated launch, compared with launching the satellites individually. Most significantly, the launch cost when using carrier vehicles is primarily determined by the total number of satellites in the constellation, rather than the number of orbital planes. Thus, the carrier vehicle deployment strategy would allow for constellations with a large number of planes to be deployed for a fraction of the equivalent cost if traditional launch methods were used

Mapping the Nation : Towards a National Land-Use Map for Scotland

[Climate change; urban sprawl; habitat destruction. Our world is changing – and we need to start watching. Scottish organisations are trying to monitor and manage our rapidly changing world. Scotland is also at the centre of Europe’s “New Space” revolution, with more satellites built in Glasgow than anywhere else in Europe. This project aims to create a roadmap to harness this abundant space data for a Scottish National Land-Use Map, to better monitor, manage, and protect our rapidly changing world.

Applications of responsive small satellites with MIT TILE electrospray propulsion

Responsive, manoeuvrable small satellites are an enabling technology for affordable, flexible and agile space missions with possible applications as wide-reaching as military reconnaissance, disaster response, and even wildlife tracking. This paper presents an analysis of some of these applications and is the outcome of a four month collaborative research visit at the Space Propulsion Laboratory of the Massachusetts Institute of Technology. This work builds upon the analytical satellite manoeuvring strategy previously developed by the author, and analyses the potential capabilities and applications of small satellites equipped with the MIT TILE electrospray thruster. This previously developed analytical method enables the rapid investigation of the manoeuvres of a constellation of small satellites, with the goal of targeting a particular region on the Earth. A full overview of the solution space can be rapidly generated, allowing for the mission designer or operator to trade off all possible manoeuvres and select the best solution for their specific purpose. The MIT TILE is a modular, miniaturised MEMS based propulsion system for nanosatellites capable of producing 350μN nominal thrust for up to 200hrs operation. A standard TILE system weighs <450g and is sized to fit in 0.5U of a CubeSat. Three case studies are presented which demonstrate the effectiveness of responsive satellites in disaster response missions. The first case study considers a rapid flyover of Los Angeles following an earthquake. The results show a reduction in flyover time of almost 9 days using 21m/s Δ퐕 when compared with a non-manoeuvring satellite. A second case study considers a fire detection constellation of 24 satellites, which can manoeuvre to provide targeted coverage of a given region as required. Selecting the Cairngorms National Park in Scotland, UK as the region of interest, the results show that by manoeuvring the constellation to directly target the region, an increase in coverage is achievable over the entire target area, with total coverage times of some areas more than doubled from 3.4 minutes coverage in a week to 8.4 minutes. The final case study considers providing communication services to helicopters at a range of locations from the UK to Svalbard, Norway. The manoeuvring capabilities of the satellites are used to follow the helicopters over an eight week period. Results show that a single satellite using <150m/s ΔV can achieve 50 flyovers of the helicopters during the journey, compared with 24 flyovers if a static satellite is used. These missions are all shown to be possible with existing technologies, and they exemplify the dramatic improvement in performance that can be achieved by using manoeuvrable satellites

General perturbation method for satellite constellation reconfiguration using low-thrust maneuvers

A general perturbation solution to a restricted low-thrust Lambert rendezvous problem, considering circular-to-circular in-plane maneuvers using tangential thrust and including a coast arc, is developed. This provides a fully analytical solution to the satellite reconnaissance problem. The solution requires no iteration. Its speed and simplicity allow problems involving numerous spacecraft and maneuvers to be studied; this is demonstrated through two case studies. In the first, a range of maneuvers providing a rapid flyover of Los Angeles is generated, giving an insight to the trade space and allowing the maneuver that best fulfills the mission to be selected. A reduction in flyover time from 13.8 to 1.6 days is possible using a less than 17 m∕s velocity change. A comparison with a numerical propagator including atmospheric friction and an 18th-order tesseral model shows 4 s of difference in the time of flyover. A second study considers a constellation of 24 satellites that can maneuver into repeating ground track orbits to provide persistent coverage of a region. A set of maneuvers for all satellites is generated for four sequential targets, allowing the most suitable maneuver strategy to be selected. Improvements in coverage of greater than 10 times are possible as compared to a static constellation using 35% of the propellant available across the constellation

Requirements for a global lidar system: spaceborne lidar with wall-to-wall coverage

Lidar is the optimum technology for measuring bare Earth elevation beneath, and the structure of, vegetation. Consequently airborne laser scanning (ALS) is widely employed for use in a wide range of applications. However, ALS is not available globally nor frequently updated due to its high cost per unit area. Spaceborne lidar can map globally, but energy requirements limit existing spaceborne lidars to sparse sampling missions unsuitable for many common ALS applications. This paper derives the equations to calculate the coverage a lidar satellite could achieve for a given set of characteristics (and released open-source), then uses a cloud map to determine the number of satellites needed to achieve continuous, global coverage within a certain time-frame. Using the characteristics of existing in-orbit technology, a single lidar satellite could have a continuous swath width of 300 m when producing a 30 m resolution map. Consequently 12 satellites would be needed to produce a continuous map every five years, increasing to 418 satellites for 5 m resolution. Building twelve of the currently in-orbit lidar systems is likely to be prohibitively expensive and so the potential of technological developments to lower the cost of a GLS are discussed. Once these technologies achieve a sufficient readiness level, a Global Lidar System could be cost-effectively realised

Design of main propulsion system for a reusable suborbital rocket

In recent years there has been an increased interest in the use of CubeSats to perform research in the realms of microgravity and earth observation. Previously, CubeSats have generally been placed into orbit as secondary payloads, piggy-backing on the launches of larger spacecraft. This has meant that CubeSat orbits and launch schedules have been decided by the requirements of other missions, restricting the manner in which they can be used. Due to the lack of flexibility in mission design afforded by traditional launch options, and the increasing competition for CubeSat launch spots, it has become desirable to develop a dedicated small satellite launch platform. This would allow for the execution of more novel and exciting missions, utilising orbits specifically designed with small satellites in mind. Tranquility Aerospace Ltd are currently engaged in the design of a two-stage vertical take-off and landing (VTVL) launcher, aimed at the small satellite market. Due to the many engineering challenges involved, they are aiming to first develop a suborbital launch vehicle in order to test the technologies necessary. This launch vehicle will be single-stage, and capable of vertical take-off and landing. As a student project at the University of Strathclyde, the main rocket propulsion system for this vehicle is being designed. This paper will outline the development of the propulsion system, including the propellant feed system, injector plate, thrust chamber and thermal control system. The key design driver is to lower the overall system mass, including the mass of the propellant. Comparisons of the impact of different subsystem configurations on performance will be assessed and discussed, and a focus will be placed on aspects of the design impacted by the requirement for reusability. The goal is to produce a fully workable design which is ready for manufacture and can be taken forward to the testing phase of development

Investigation of Very Low Earth Orbits (VLEOs) for global spaceborne Lidar

Very Low Earth Orbits (VLEOs) have been proposed as a beneficial space mission regime due to their propensity to increase instrument spatial resolution and reduce launch cost per unit mass. However, for visual instruments, these benefits come at the cost of a decreased instrument swath width. This reduction results in longer revisit periods for regions on Earth and longer time until global coverage is achieved. Conversely, light detection and ranging (lidar) as an active remote sensing technique, can benefit from larger swath widths at lower altitudes, due to the increased signal-to-noise ratio. Investigation of this relationship shows that lidar swath width is inversely proportional to altitude squared, and, as a result, the number of spacecraft required to provide a desired lidar coverage also decreases approximately in inverse proportion to altitude squared. Investigation of suitable propulsion systems shows that although propellant mass and number of thrusters required for orbit maintenance increases with decreasing altitude, the overall system mass, and hence launch cost, will, in general, tend to decrease with decreasing altitude due to the lower number of spacecraft required. For a given mission, spacecraft bus, and propulsion system, a VLEO altitude can be identified that will result in the minimum overall mission cost