Near Earth Objects (NEOs) are comets or asteroids that intersect or pass near to our planet posing a real and underestimated danger to mankind. While the probability of impact is low, the consequences of such an impact could be apocalyptic. Various programs are underway to discover these kilometer-sized objects from Earth. However, once targets of interest have been identified a fly-by or orbiting spacecraft is required to understand the objects' mass, morphology and composition.Fly-past NEO missions represent the simplest interplanetary missions and need not be high cost. An 'entry level' mission has been conceptually designed able to deliver a 10kg science 'reference' payload to NEO fly-by for a total mission cost (including launch and operations) of Euro20million (FY2003), This paper outlines the platform architecture, cost and cost drivers, and describes the key technology trades to be performed and the developments required to extend current Low Earth Orbit (LEO) technology to a deep space mission. It concludes by identifying the top-level trade-offs to be made in order to enhance the science return of the mission.

The Venus Entry Probe study is one of ESA's technology reference studies. It aims to identify; the technologies required to develop a low-cost, science-driven mission for in-situ exploration of the atmosphere of Venus, and the philosophy that can be adopted. The mission includes a science gathering spacecraft in an elliptical polar Venus orbit, a relay satellite in highly elliptical Venus orbit, and an atmospheric entry probe delivering a long duration aerobot which will drop several microprobes during its operational phase. The atmospheric entry sequence is initiated at 120 km altitude and an entry velocity of 9.8 kms. Once the velocity has reduced to 15 ms the aerobot is deployed. This consists of a gondola and balloon and has a floating mass of 32 kg (which includes 8 kg of science instruments and microprobes). To avoid Venus’ crushing surface pressure and high temperature an equilibrium float altitude of around 55 km has been baselined. The aerobot will circumnavigate Venus several times over a 22-day period analysing the Venusian middle cloud layer. Science data will be returned at 2.5 kbps over the mission duration. At scientifically interesting locations 15 drop-sondes will be released. This paper focuses on the final mission design with particular emphasis on system level trade-offs including the balloon and pressurisation system, communications architecture, power system, design for mission lifetime in a hostile and acidic environment. It discusses the system design, design drivers and presents an overview of the innovative missionenabling and mission-enhancing technologies. Introduction The Venus Entry Probe is one of ESA’s Technology Reference Studies (TRS). These are model sciencedriven missions that although not part of the ESA science programme are able to provide focus to future technology requirements. This is accomplished through the study of several technologically demanding and scientifically meaningful mission concepts, which are strategically chosen to address diverse technological issues. The TRSs complement ESA’s current mission specific development programme and allow the ESA Science Directorate to strategically plan the development of technologies that will enable potential future scientific missions. Key technological objectives for future planetary exploration include the use of small orbiters and in-situ probes with highly miniaturized and highly integrated payload suites. These low resource, and therefore potentially low cost, spacecraft allow for a phased strategic approach to planetary exploration, thus reducing mission risks compared to a single heavy resource mission. The aim of the Venus Entry Probe (VEP) TRS is to study approaches for low cost in-situ exploration of the Venusian atmosphere. The mission profile consists of two minisatellites, one dedicated to atmospheric remote sensing and the other specialised for entry probe deployment as well as data relay . This two-satellite configuration is required in order to commence the remote sensing atmospheric investigations prior to the aerobot deployment. The additional advantage is that through the use of a data relay satellite, the other minisatellite can practically continuously perform remote sensing investigations of the atmosphere. The Low Venus Orbiter (LVO) enters low Venus orbit (6000km x 2000km) and contains a highly integrated remote sensing payload suite primarily dedicated to support the in-situ atmospheric measurements of the aerobot and to address the global atmospheric science objectives. The Venus Relay Satellite (VRS) enters a highly elliptical orbit (215,000 x 250 km), deploys the Venus Entry Vehicle (VEV) and subsequently operates as a data relay satellite (and may provide navigational support). The aerobot consists of a long-duration balloon and gondola (depicted below) that will analyse the Venusian middle Venusian middle cloud layer at an altitude of ~55 km, where the environment is relatively benign. The balloon will deploy a swarm of active ‘ballast’ micro-sondes, which, once deployed, will determine vertical profiles of the lower atmosphere . Figure 1 Venus aerobot mission (gondola and balloon) Mission Objectives The objective of the Venus Entry Probe TRS is to establish a feasible mission profile for a low-cost insitu exploration of Venus. The primary scientific objectives of the mission are to study: 1. Origin and evolution of the atmosphere 2. Composition and chemistry of the lower atmosphere 3. Atmospheric dynamics 4. Aerosols in the cloud layers A more detailed description of the scientific rationale is detailed by . The strategy for this mission development is to meet the science requirements at lowest overall mission cost. The study will determine the mission cost, the system drivers and determine if the instrument duty cycle is viable. It will also identify technologies required to develop such a mission. Mission Design MISSION REQUIREMENTS In order to address the science objectives, the following mission requirements have been imposed on the Venus Entry Vehicle: Mission launch in 2014 onwards Planetary protection requirements: None Support ~4 kg payload suite as well as ~4 kg microprobes, including a ranging and navigation system (DALOMIS-C). This generates science data at a rate of 2.5kbs for the duration of the mission Deploy swarm of fifteen 115 g drop sondes or microprobes. These will either be deployed individually or in groups of 3 in a drop campaign Nominal mission duration: 15 days Extended mission duration: to 30 days Ballistic or orbital entry is permitted, entry must be 20±5 latitude either north or south Maximum entry deceleration is ~200 G Entry sequencing must be dual redundant Aerobot float altitude extremes are 53-62 km. For the first 8 days the balloon must be at an equilibrium float altitude of 55 km

Traditionally interplanetary spacecraft are expensive, and have development schedules lasting up to 10 years. Being so high cost few concepts reach the manufacture phase. As a result of the infrequency of these missions, more experiments and supporting systems become incorporated, further driving up the cost, and delaying the schedule. Programmes rapidly become unaffordable. The Venus Entry Probe is one of ESA's technology reference studies. It aims to identify the technologies required to develop a low cost science-driven mission for in-situ exploration of the atmosphere of Venus, and the philosophy which can be adopted. The mission includes a science gathering spacecraft in circular polar Venus orbit, a relay satellite in elliptical Venus orbit, and an atmospheric entry probe delivering a long duration aerobot, which will drop several microprobes during its operational phase. This paper will focus on the design of the mission, spacecraft and aerobot, with particular emphasis on system level trade-offs. Recognising that resource reduction (mass, power, volume etc.) will be a key requirement to achieving low recurring cost spacecraft, the project has identified a variety of innovative mission enabling and mission enhancing technologies.

One of the potentially biggest pitfalls academic institutions can fall into when attempting to design new microsatellite missions is to start from scratch. A completely open-ended problem creates a nearly infinite time sink for students and faculty as they struggle to meet primary mission requirements while inventing the overall system architecture. Fortunately, commercial-off-the-shelf spacecraft architectures offer the opportunity to significantly bound the design space, allowing students and faculty to devote scarce time and other resources to the most important aspects of mission design and overall systems engineering. This paper describes recent experience at the USAF academy in applying the "SNAP" nanosatellite architecture, designed by Surrey Satellite Technology, UK, to the FalconSAT-2 program. The SNAP architecture is first described followed by a discussion of its application to FalconSAT-2, including how it influenced system design decisions and accelerated system development.

The capture effect can improve throughput of random multiple access protocols. An exact model for evaluating the capture effect is proposed for noncoherent FSK demodulation (NCFSK) under a slow fading assumption. This model takes into account the bit decision process of the NCFSK demodulator in the presence of both co-channel interference and noise. An approximate and simplified model is also presented for comparison. Additionally, a Gaussian assumption-based model is derived, which turns out to be too pessimistic for slow fading channels. The capture effect is investigated by means of the proposed exact model in a low Earth orbit satellite channel. The result shows that a relatively large link margin can result in a large capture probability, and consequently benefit slotted ALOHA protocols with a substantial throughput improvement.

In order to meet the growing global requirement for affordable missions beyond Low Earth Orbit, two types of platform are under design at the Surrey Space Centre. The first platform is a derivative of Surrey's UoSAT-12 minisatellite, launched in April 1999 and operating successfully in-orbit. The minisatellite has been modified to accommodate a propulsion system capable of delivering up to 1700 m/s delta-V, enabling it to support a wide range of very low cost missions to LaGrange points, Near-Earth Objects, and the Moon. A mission to the Moon - dubbed “MoonShine” - is proposed as the first demonstration of the modified minisatellite beyond LEO. The second platform - Surrey's Interplanetary Platform - has been designed to support missions with delta-V requirements up to 3200 m/s, making it ideal for low cost missions to Mars and Venus, as well as Near Earth Objects (NEOs) and other interplanetary trajectories. Analysis has proved mission feasibility, identifying key challenges in both missions for developing cost-effective techniques for: spacecraft propulsion; navigation; autonomous operations; and a reliable safe mode strategy. To reduce mission risk, inherently failure resistant lunar and interplanetary trajectories are under study. In order to significantly reduce cost and increase reliability, both platforms can communicate with low-cost ground stations and exploit Surrey's experience in autonomous operations. The lunar minisatellite can provide up to 70 kg payload margin in lunar orbit for a total mission cost US$16–25 M. The interplanetary platform can deliver 20 kg of scientific payload to Mars or Venus orbit for a mission cost US$25–50 M. Together, the platforms will enable regular flight of payloads to the Moon and interplanetary space at unprecedented low cost. This paper outlines key systems engineering issues for the proposed Lunar Minisatellite and interplanetary Platform Missions, and describes the accommodation and performance offered to planetary payloads.

Capture effect can improve the throughput of random multiple access protocols. An evaluation model of the capture effect is proposed for coherent BPSK demodulation. The model introduces the concept of carrier synchronisation probability conditioned upon the carrier-to-interference ratio into the evaluation of the capture probability. Moreover, the effect of bit-sync offset of interfering packets on the capture probability is also taken into account. An attempt is made by means of the proposed model to investigate the capture effect in the LEO satellite channel. The result shows that a relatively large link margin can result in a remarkable capture probability, and consequently benefit slotted ALOHA protocols with a substantial throughput improvement.

To efficiently support bursty short message transmission in little LEO satellite communication networks, ALOHA protocols with ACK-list and event-based slot timing are proposed. The capture effect, commonly existing in fading channels, is exploited to improve the throughput of the protocol. This paper makes the first attempt to investigate the capture effect in LEO satellite channels, and outlines a picture of the capture effect. Moreover, a novel scheme to enhance the capture effect by a protocol approach is proposed. The simulation results show that the division of a footprint into ring-shaped areas, in combination with a biased transmitting power assignment substantially strengthens the capture effect, and consequently improves the throughput of the protocol

The capture effect can improve the channel throughput of random multiple access protocols. So far, no attempt to investigate the capture effect in LEO satellite channels has been made. Two sophisticated evaluation models are proposed respectively for coherent BPSK and non-coherent FSK demodulation. The C-BPSK model introduces the threshold of carrier lock into the calculation of the capture probability, and consequently removes the overestimated part of the probability. Moreover, the affect of bit-sync offsets of the interfering packets on the capture probability is also taken into account. The novel NC-FSK model properly considers the bit decision process of the receiver under interfering packets, and is able to present a more practical estimate of the capture effect in the UoSAT satellite channel. Although originally proposed for the LEO satellite channel, the two models are completely applied to the other mobile channels where the channel noise is Gauss-distributed.

Earth remote sensing (alongside communications) is one of the key application of Earth-orbiting satellites. Civilian satellites in the LANDSAT and SPOT series provide Earth images which have been used for a vast spectrum of applications in agriculture, meteorology, hydrology, urban planning and geology, to name but a few. In the defence sector, satellite remote sensing systems are a critical tool in strategic and tactical planning – for the countries which can afford them. To date, remote sensing satellites have fallen into one of these two categories: military missions driven by the requirement for very high resolution and orbital agility; and multipurpose civil satellites using general purpose sensors to serve a diverse community of end users. For military-style missions, the drive to high resolution sets the requirements for optics, attitude control and downlink data bandwidth. For civil missions, the requirement to satisfy multiple, diverse user applications forces compromises on spectral band and orbit selection. Although there are exceptions, many small satellite remote sensing missions carry on in this tradition, concentrating on ultra high resolution products for multiple user communities. This results in satellites costing on the order of US $100 M, not optimised for any particular application. This paper explores an alternative path to satellite remote sensing, aiming simultaneously to reduce cost and to optimise imaging products for specific applications. By decreasing the cost of the remote sensing satellite system to a critical point, it becomes appropriate to optimise the sensor's spectral and temporal characteristics to fit the requirements of a small, specialised user base. The critical engineering trade-off faced in a cost driven mission is how to reduce mission cost while still delivering a useful product to the selected user. At the Surrey Space Centre, we have pursued an engineering path using two dimensional CCD array sensors, commercial off-the-shelf lenses and gravity-gradient stabilised microsatellites. In spite of the inherent limitations of such systems, recent successes with the Thai Microsatellite Company's Thai-Phutt satellite show that a system costing in the region of US $3 million, can approach the spectral and spatial characteristics of LANDSAT. Surrey's UoSAT-12 minisatellite (to be launched April, 1999) will further develop this cost-driven approach to provide 10 m panchromatic resolution and 30 m multi-spectral resolution. This paper describes the Thai-Phutt and UoSAT-12 imaging systems, explaining the engineering methods and trade-offs. Although Surrey is presently the only centre presently pursuing such implementations, our paper shows that they deserve wider consideration.

The unique characteristics of a LEO satellite global email communication network have been identified. A new type of network system architecture optimized for the LEO satellite email network is proposed, in which the LEO satellite is designed as an active switching node using a specifically designed LEO satellite network addressing mechanism. New LEO channel optimization algorithms are also proposed to better use the up/down link resource. Comparing to the current design of the system in which the satellite works as a mailbox, a fifty percent performance improvement can be expected.