X-ray, γ-ray and neutron detector development for future space instrumentation

An overview of the material developments currently being pursued by the European Space Agency (ESA) with respect to X-ray, γ-ray and neutron detection media is given. The benefits underlying the selection of specific materials are discussed. In the X-ray regime the basic microphysics behind degraded spectral performance caused by the polarization effect and Te inclusions, have been investigated. We have found that both effects degrade the spectra by producing a localised distortion of the electric field. The underlying cause for the distortion does however differ. For the γ-ray regime, a low-noise equivalent scintillator to LaBr₃ has been developed in the form of CeBr₃. Three inch crystals with a FWHM energy resolution of 4.4% at 662 keV have successfully been produced. For neutron detection, a boron compound based solid state neutron detector has been developed that is sensitive to neutrons, alpha particles and 60 keV X-rays. The measured detection efficiency in detection for thermal neutrons was found to be in the order of a few per cent.     

On the capabilities of nano electrokinetic thrusters for space propulsion

A theoretical analysis considering the capabilities of nano electrokinetic thrusters for space propulsion is presented. The work describes an electro-hydro-dynamic model of the electrokinetic flow in nano-channels and represents the first attempt to exploit the advantages of the electrokinetic effect as the basis for a new class of nano-scale thrusters suitable for space propulsion. Among such advantages are their small volume, fundamental simplicity, overall low mass, and actuation efficiency. Their electrokinetic efficiency is affected by the slip length, surface charge, pH and molarity. These design variables are analyzed and optimized for the highest electrokinetic performance inside nano-channels. The optimization is done for power consumption, thrust and specific impulse resulting in high theoretical efficiency ∼99% with corresponding high thrust-to-power ratios. Performance curves are obtained for the electrokinetic design variables showing that high molarity electrolytes lead to high thrust and specific impulse values, whereas low molarities provide highest thrust-to-power ratios and efficiencies. A theoretically designed 100 nm wide by 1 μm long emitter optimized using the ideal performance charts developed would deliver thrusts from 5 to 43 μN, specific impulse from 60 to 210 s, and would have power consumption between 1–15 mW. It should be noted that although this is a detail analytical analysis no prototypes exist and any future experimental work will face challenges that could affect the final performance. By designing an array composed of thousands of these single electrokinetic emitters, it would result in a flexible and scalable propulsion system capable of providing a wide range of thrust control for different mission scenarios and maintaining very high efficiencies and thrust-to-power ratio by varying the number of emitters in use at any one time.     

Venus surface power and cooling systems

A mission to the surface of Venus would have high scientific value, but most electronic devices and sensors cannot operate at the 450 °C ambient surface temperature of Venus. Power and cooling systems were analyzed for Venus surface operation. A radioisotope power and cooling system was designed to provide electrical power for a probe operating on the surface of Venus. For a mission duration of substantial length, the use of thermal mass to maintain an operable temperature range is likely impractical, and active refrigeration may be required to keep components at a temperature below ambient. Due to the high thermal convection of the high-density atmosphere, the heat rejection temperature was assumed to be at a 500 °C radiator temperature, 50 °C above ambient. The radioisotope Stirling power converter designed produces a thermodynamic power output capacity of 478.1 W, with a cooling power of 100 W. The overall efficiency is calculated to be 23.36%. The mass of the power converter is estimated at approximately 21.6 kg.     

BBM/EM design of the thermal hyperspectral imager: An instrument for remote sensing of earth's surface,atmosphere and ocean,from a microsatellite platform

The Thermal Hyperspectral Imager (THI) is a low cost, low mass, power efficient instrument designed to acquire hyperspectral remote sensing data in the long-wave infrared. The instrument has been designed to satisfy mass, volume, and power constraints necessary to allow for its accommodation in a 95 kg micro-satellite bus, designed by staff and students at the University of Hawai'i. THI acquires approximately 30 separate spectral bands in the 8–14 μm wavelength region, at 16 wavenumber resolution. Rather than using filtering or dispersion to generate the spectral information, THI uses an interferometric technique. Light from the scene is focused onto an uncooled microbolometer detector array through a stationary interferometer, causing the light incident at each detector at any instant in time to be phase shifted by an optical path difference which varies linearly across the array in the along-track dimension. As platform motion translates the detector array in the along-track direction at a rate of approximately one pixel per frame (the camera acquires data at 30 Hz) the radiance from each scene element can be sampled at each OPD, thus generating an interferogram. Spectral radiance as a function of wavelength is subsequently obtained for each scene element using standard Fourier transform techniques. Housed in a pressure vessel to shield COTS parts from the space environment, the total instrument has a mass of 15 kg. Peak power consumption, largely associated with the calibration procedure, is <90 W. From a nominal altitude of 550 km the resulting data would have a spatial resolution of approximately 300 m. Although an individual imaging event yields approximately 1 Gbit of raw uncompressed data, onboard processing (to convert the interferograms into a conventional spectral hypercube) can reduce this to tens of Mega bits per scene. In this presentation we will describe (a) the rationale for the project, (b) the instrument design, and (c) how the data are processed. Finally we will present data acquired by THI on a laboratory microscope stage to demonstrate the spectro-radiometric quality of the data that the instrument can provide.     

Launch and early operations of Herschel and Planck

On 14 May 2009 the European Space Agency launched 2 space observatories: Herschel (with a 3.5 m mirror it is the largest space telescope ever) will collect long-wavelength infrared radiation and will be the only space observatory to cover the spectral range from far-infrared to sub-millimetre wavelengths, and Planck will look back at the dawn of time, close to the Big Bang, and will examine the Cosmic Microwave Background (CMB) radiation to a sensitivity, angular resolution and frequency range never achieved before. This paper will present the Flight Dynamics, mission analysis challenges and flight results from the first 3 months of these missions.Both satellites were launched on the same Ariane 5 and travelled to the L2 Lagrange point of the sun–earth system 1.5 million km from the earth in the opposite direction of the sun. There they were injected to a quasi-halo orbit (Herschel) with the dimension of typically 750,000 km×450,000 km, and a Lissajous orbit (Planck) of 300,000 km×300,000 km.In order to reach these Lissajous orbits it is mandatory to perform large trajectory correction manoeuvres during the first days of the mission. Herschel had its main manoeuvres on the first day. Planck had to be navigated on the first day and by a mid-course correction manoeuvre, the L2 orbit insertion manoeuvre was planned on day 50. If these slots were missed, fuel penalties would rapidly increase.This posed a heavy load on the operations teams because both spacecrafts have to be thoroughly checked out and put into the correct modes of their attitude control systems during the first hours after launch.The sequence of events will be presented and explained and the orbit determination results as well as the manoeuvre planning will be emphasised.     

Modelling of the optical observations of a geosynchronous fragmentation event

A study of the evolution and optical detectability of a fragmentation debris cloud in geosynchronous orbit has been carried out. The 1998 NASA breakup model has been used to generate orbit data for 95 fragments larger than 10 cm size from a 1000 kg satellite. The orbital evolution of these fragments is studied using a precision numerical propagator, employing a high-fidelity force model. Although the fragments rapidly disperse throughout the geostationary arc, they remain localised in right ascension of ascending node and inclination, and are driven along a narrow inertial corridor by luni-solar perturbations. The ESA PROOF software is used to study the detectability of the fragments using a 1- and 0.5-m telescope design. The 1-m telescope can detect 82% of the fragments (down to 13 cm in size) whilst the 0.5-m telescope can detect 39% of the fragments (down to 30 cm size). Due to the large along-track spread of the fragments, a time limit of 1-month post-breakup can be established for a space surveillance system to catalogue the breakup fragments. After this time the angular separation is such that the fragments disperse into the background population, and are no longer distinguishable as originating from a common breakup event.     

Response of Phaseolus vulgaris L. plants to low-let ionizing radiation: Growth and oxidative stress

The scenarios for the long-term habitation of space platforms and planetary stations involve plants as fundamental part of Bioregenerative Life Support Systems (BLSS) to support the crew needs. Several constraints may limit plant growth in space: among them ionizing radiation is recognized to severely affect plant cell at morphological, physiological and biochemical level. In this work, plants of Phaseolus vulgaris L. were subjected to four different doses of X-rays (0.3, 10, 50 and 100 Gy) in order to assess the effects of ionizing radiation on this species and to analyze possible mechanisms carried out to overcome the radiation injuries. The effects of X-rays on plant growth were assessed by measuring stem elongation, number of internodes and leaf dry weight. The integrity of photosynthetic apparatus was evaluated by photosynthetic pigment composition and ribulose 1,5-bisphosphate carboxylase (Rubisco) activity, whereas changes in total antioxidant pool and glutathione S transferase activity (GST) were utilized as markers of oxidative stress. The distribution of phenolic compounds in leaf tissues as natural shielding against radiation was also determined.Irradiation of plants at 0.3 and 10 Gy did not determine differences in all considered parameters as compared to control. On the contrary, at 50 and 100 Gy a reduction of plant growth and a decrease in photosynthetic pigment content, as well as an increase in phenolic compounds and a decrease in total antioxidant content and GST activity were found. Only a slight reduction of Rubisco activity in leaves irradiated at 50 and 100 Gy was found. The overall results indicate P. vulgaris as a species with a good potential to face ionizing radiation and suggest its suitability for utilization in BLSSs.     

Sleep and cognitive function of crewmembers and mission controllers working 24-h shifts during a simulated 105-day spaceflight mission

The success of long-duration space missions depends on the ability of crewmembers and mission support specialists to be alert and maintain high levels of cognitive function while operating complex, technical equipment. We examined sleep, nocturnal melatonin levels and cognitive function of crewmembers and the sleep and cognitive function of mission controllers who participated in a high-fidelity 105-day simulated spaceflight mission at the Institute of Biomedical Problems (Moscow). Crewmembers were required to perform daily mission duties and work one 24-h extended duration work shift every sixth day. Mission controllers nominally worked 24-h extended duration shifts. Supplemental lighting was provided to crewmembers and mission controllers. Participants' sleep was estimated by wrist-actigraphy recordings. Overall, results show that crewmembers and mission controllers obtained inadequate sleep and exhibited impaired cognitive function, despite countermeasure use, while working extended duration shifts. Crewmembers averaged 7.04±0.92 h (mean±SD) and 6.94±1.08 h (mean±SD) in the two workdays prior to the extended duration shifts, 1.88±0.40 h (mean±SD) during the 24-h work shift, and then slept 10.18±0.96 h (mean±SD) the day after the night shift. Although supplemental light was provided, crewmembers’ average nocturnal melatonin levels remained elevated during extended 24-h work shifts. Naps and caffeine use were reported by crewmembers during ∼86% and 45% of extended night work shifts, respectively. Even with reported use of wake-promoting countermeasures, significant impairments in cognitive function were observed. Mission controllers slept 5.63±0.95 h (mean±SD) the night prior to their extended duration work shift. On an average, 89% of night shifts included naps with mission controllers sleeping an average of 3.4±1.0 h (mean±SD) during the 24-h extended duration work shift. Mission controllers also showed impaired cognitive function during extended duration work shifts.These findings indicate that extended duration work shifts present a significant challenge to crewmembers and mission support specialists during long-duration space mission operations. Future research is needed to evaluate the efficacy of alternative work schedules and the development and implementation of more effective countermeasures will be required to maintain high levels of performance.     

Co-deposition of Ir-containing Zr coating by double glow plasma

Ir–Zr co-deposition coatings with 71 at% Zr were deposited on graphite by double glow plasma at 1073–1123 K. The structure and composition of the coatings were confirmed by FE-SEM, XRD, XPS and EDS. The hardness and the elastic modulus of the coatings were estimated by nanoindentation instrument. The adhesion strength between the coating and the substrate was evaluated by a scratch tester. The results showed that the coating was composed of nanocrystalline grains with a size of 80–90 nm compared with 0.5 μm for the pure Ir coating. The fine grains of the coatings might be attributed to the additional Zr element. New phases IrZr and ZrC were formed due to the high content of Zr and high deposition temperature. The hardness and elastic modulus of the coatings were about 7.5 GPa and 388 GPa, respectively. The adhesive force between the coating and the substrate was about 10 N.     

Radiation engineering analysis of shielding materials to assess their ability to protect astronauts in deep space from energetic particle radiation

An analysis is performed on four typical materials (aluminum, liquid hydrogen, polyethylene, and water) to assess their impact on the length of time an astronaut can stay in deep space and not exceed a design basis radiation exposure of 150 mSv. A large number of heavy lift launches of pure shielding mass are needed to enable long duration, deep space missions to keep astronauts at or below the exposure value with shielding provided by the vehicle. Therefore, vehicle mass using the assumptions in the paper cannot be the sole shielding mechanism for long duration, deep space missions. As an example, to enable the Mars Design Reference Mission 5.0 with a 400 day transit to and from Mars, not including the 500 day stay on the surface, a minimum of 24 heavy lift launches of polyethylene at 89,375 lbm (40.54 tonnes) each are needed for the 1977 galactic cosmic ray environment. With the assumptions used in this paper, a single heavy lift launch of water or polyethylene can protect astronauts for a 130 day mission before exceeding the exposure value. Liquid hydrogen can only protect the astronauts for 160 days. Even a single launch of pure shielding material cannot protect an astronaut in deep space for more than 180 days using the assumptions adopted in the analysis. It is shown that liquid hydrogen is not the best shielding material for the same mass as polyethylene for missions that last longer than 225 days.     

Deployment history and design considerations for space reactor power systems

The history of the deployment of nuclear reactors in Earth orbits is reviewed with emphases on lessons learned and the operation and safety experiences. The former Soviet Union's “BUK” power systems, with SiGe thermoelectric conversion and fast neutron energy spectrum reactors, powered a total of 31 Radar Ocean Reconnaissance Satellites (RORSATs) from 1970 to 1988 in 260 km orbit. Two of the former Soviet Union's TOPAZ reactors, with in-core thermionic conversion and epithermal neutron energy spectrum, powered two Cosmos missions launched in 1987 in ～800 km orbit. The US’ SNAP-10A system, with SiGe energy conversion and a thermal neutron energy spectrum reactor, was launched in 1965 in 1300 km orbit. The three reactor systems used liquid NaK-78 coolant, stainless steel structure and highly enriched uranium fuel (90–96 wt%) and operated at a reactor exit temperature of 833–973 K. The BUK reactors used U-Mo fuel rods, TOPAZ used UO₂ fuel rods and four ZrH moderator disks, and the SNAP-10A used moderated U-ZrH fuel rods. These low power space reactor systems were designed for short missions (～0.5 kW_e and ～1 year for SNAP-10A, <3.0 kW_e and <6 months for BUK, and ～5.5 kW_e and up to 1 year for TOPAZ). The deactivated BUK reactors at the end of mission, which varied in duration from a few hours to ～4.5 months, were boosted into ～800 km storage orbit with a decay life of more than 600 year. The ejection of the last 16 BUK reactor fuel cores caused significant contamination of Earth orbits with NaK droplets that varied in sizes from a few microns to 5 cm. Power systems to enhance or enable future interplanetary exploration, in-situ resources utilization on Mars and the Moon, and civilian missions in 1000–3000 km orbits would generate significantly more power of 10's to 100's kW_e for 5–10 years, or even longer. A number of design options to enhance the operation reliability and safety of these high power space reactor power systems are presented and discussed.     

Super-capacitor energy storage for micro-satellites: Feasibility and potential mission applications

Small satellites, weighting between 100 and 200 kg, have witnessed increasing use for a variety of space applications including remote sensing constellations and technology demonstrations. The energy storage/stored power demands of most spacecraft, including small satellites, are currently accommodated by rechargeable batteries—typically nickel–cadmium cells (specific energy of 50 Wh kg⁻¹), or more recently lithium-ion cells (150 Wh kg⁻¹). High energy density is a primary concern for spacecraft energy storage design, and these batteries have been sufficient for most applications. However, constraints on the allowable on-board battery size have limited peak power performance such that the maximum power supply capability of small satellites currently ranges between only 70 and 200 W. This relatively low maximum power limits the capabilities of small satellites in terms of payload design and selection. In order to enhance these satellites' power performance, the research reported in this paper focused on the implementation of super-capacitors as practical rechargeable energy storage medium, and as an alternative to chemical batteries. Compared to batteries, some super-capacitors are able to supply high power at high energy-efficiency, but unfortunately they still have a very low energy density (5–30 Wh kg⁻¹). However, the provision of this high power capability would considerably widen the range of small satellite applications.     

Rosetta visits asteroid (21)Lutetia

The International Rosetta Mission, cornerstone of the European Space Agency Scientific Programme, was launched on 2nd March 2004 to its 10 years journey to comet Churyumov–Gerasimenko. Rosetta will reach the comet in summer 2014, orbit it for about 1.5 years down to distances of a few Kilometres and deliver the Lander Philae onto its surface. After its successful asteroid fly-by in September 2008, Rosetta came back to Earth, for the final gravity acceleration towards its longest heliocentric orbit, up to a distance of 5.3 AU. It is during this phase that Rosetta crossed for the second time the main asteroids belt and performed a close encounter with asteroid (21)Lutetia on the 10th of July 2010 at a distance of ca. 3160 km and a relative velocity of 15 km/s. The payload complement of the spacecraft was activated to perform highly valuable scientific observations. The approach phase to the celestial body required a careful and accurate optical navigation campaign that will prove to be useful also for the comet approach phase. The experience gained with first asteroid flyby in 2008 was fed back into the operations definition and preparation for this highly critical phase; this concerns in particular the operations of the navigation camera for the close-loop autonomous asteroid tracking and of the main scientific camera for high resolution imaging. It was shortly after the flyby that Rosetta became the solar-powered spacecraft to have flown furthest from the Sun (>2.72 AU). This paper presents the activities carried out and planned for the definition, preparation and implementation of the asteroid flyby mission operations, including the test campaign conducted to improve the performance of the spacecraft and payload compared to the previous flyby. The results of the flyby itself are presented, with the operations implemented, the achieved performance and the lessons learned.     

Interaction of mental and orthostatic stressors

We assessed hemodynamic responses induced by orthostatic and mental stressors, using passive head up tilt (HUT) and mental arithmetic (MA), respectively. The 15 healthy males underwent three protocols: (1) HUT alone, (2) MA in supine position and (3) MA+HUT, with sessions randomized and ≥2 weeks apart. In relation to baseline, HUT increased heart rate (HR) (+20.4±7.1 bpm; p<0.001), mean blood pressure (MBP) (+4.7±11.3 mmHg; p<0.05), diastolic blood pressure (DBP) (+6.1±11.6 mmHg; p<0.05) and total peripheral resistance (TPR) (+155±232 dyne*s/cm⁵; p<0.001) but decreased stroke volume (SV) (?33.1±13.4 ml; p<0.001) and cardiac output (CO) (?0.6±1.0 l/min; p<0.01). MA increased HR (+8.0±6.0 bpm; p<0.001), systolic blood pressure (SBP) (+9.0±7.7 mmHg; p<0.001), MBP (+10.0±6.5 mmHg; p<0.001), DBP (+9.5±7.2 mmHg; p<0.001) and CO (+0.6±0.8 l/min; p<0.01). MA+HUT increased HR (+28.8±8.4 bpm; p<0.001), SBP (+4.6±14.3 mmHg; p<0.05), MBP (+11.2±11.6 mmHg; p<0.001), DBP (+13.5±10.1 mmHg; p<0.001) and TPR (+160±199 dyne*s/cm⁵; p<0.001) but SV (?34.5±14.6 ml; p<0.001) decreased. Mental challenge during orthostatic challenge elicited greater increases in heart rate, despite similar reductions in stroke volume such as those during orthostatic stress alone. Overall, cardiac output decreases were less with combinations of mental and orthostatic challenges in comparison to orthostasis alone. This would suggest that carefully chosen mental stressors might affect orthostatic responses of people on standing up. Therefore, additional mental loading could be a useful countermeasure to alleviate the orthostatic responses of persons, particularly in those with histories of dizziness on standing up or on return to earth from the spaceflight environment of microgravity.     

Modeling the sodium potassium droplet interactions with the low earth orbit space debris environment

The NASA/JSC sodium potassium (NaK) RORSAT coolant source and propagation model has been extended to 1 mm in diameter via a size distribution, which is an inverse power law fit that has been modified to damp out in the large size regime. This function matches the observed Haystack NaK population down to diameters of about 6 mm. The extrapolated function takes the population to arbitrarily small sizes all the while retaining the mass dominance of the 1–3 cm droplets that is observed in the Haystack data. This result is physically satisfying since the mechanism of NaK ejection appears to be a nonviolent release at low relative velocities. We propose that any NaK particles smaller than about 1 mm that exist would not be due to that mechanism. Instead, we show that such a population could be the result of subsequent collisions of NaK droplets with larger resident space objects and the micrometeoroid population. Our preliminary analysis shows that collisions between these populations are likely in the time period of 1980 through present-day. Though the result of such collisions is generally unknown it is probable that some ejecta of NaK enter the low Earth orbit (LEO) environment as a result. It is these secondary NaK droplets/particles that we contend are the likely impactors noted on returned surfaces.     

Measured surface temperatures of the Hayabusa capsule during re-entry determined from ground observation

The Hayabusa sample return capsule, which contained asteroid samples, re-entered the Earth's atmosphere on June 13, 2010. An ablative carbon-phenolic thermal protection system (TPS) was used to enable a safe return for the small capsule and the containing samples. Besides a research aircraft operated by NASA with a wide range of imaging and spectrographic cameras for remote sensing of the radiation of the Hayabusa capsule during its entry flight, observation from ground based stations has been realized. We participated in the ground based observation campaign with two instruments for spectroscopic and photometric measurements aiming to detect the surface temperature and the plasma radiation in front of the re-entering capsule. The system consists in an infrared camera and a wide range miniature fibre spectrometer. The paper presents the setup, the laboratory calibration procedure, and correction for transmission. The surface temperature of the capsule reached a peak of 3250 K when the capsule was at an altitude of 55.95 km. The thermographic camera measures independently slightly higher temperature at peak heating (3308 K).     

Thrust measurement of dimethyl ether arcjet thruster

The present paper describes thrust measurement results for an arcjet thruster using Dimethyl ether (DME) as the propellant. DME is an ether compound and can be stored as a liquid due to its relatively low freezing point and preferable vapor pressure. The thruster successfully produced high-voltage mode at DME mass flow rates above 30 mg/s, whereas it yielded low-voltage mode below 30 mg/s. Thrust measurements yielded a thrust of 0.15 N and a specific impulse of 270 s at a mass flow rate of 60 mg/s with a discharge power of 1300 W. The DME arcjet thruster was comparable to a conventional one for thrust and discharge power.     

A harness for enhanced comfort and loading during treadmill exercise in space

IntroductionLocomotor and some resistance exercises in space require a gravity replacement force in order to allow 1g-like ground reaction forces to be generated. Currently bungee cords, or other loading devices, interface with the crew member through a harness with a waist belt and shoulder straps. Crew members often find the application of the required loads to be uncomfortable, particularly at the hips.MethodsAn experimental harness was built that differed from previous in-flight designs by having a wider, moldable waist belt and contoured shoulder straps with additional padding. Eight subjects ran at 100% body weight (BW) loading for a total duration of 30 min per day on 12 days over a 3-week period in simulated 0-g conditions using horizontal suspension. A 100 mm Visual Analog Scale (VAS)¹ was used to assess harness-related and lower extremity discomfort at the end of each run.ResultsThe overall rating of harness discomfort decreased from 27 mm on the 100 mm scale on day 1 to 10 mm on day 12, with significant decreases recorded for the back and hip regions as well as the overall harness.DiscussionThe experimental harness allows for repeated exposure to 30-minute bouts of 100% BW loaded simulated 0-g running with levels of discomfort less than 30 mm on a VAS scale of 0–100 mm. We believe that the use of such a harness during on-orbit exercise countermeasures may allow exercise to be performed at levels which are more effective in preventing bone and muscle loss.     

Very high delta-V missions to the edge of the solar system and beyond enabled by the dual-stage 4-grid ion thruster concept

A new and innovative type of gridded ion thruster, the “Dual-Stage 4-Grid” or DS4G concept, has been proposed and its predicted high performance validated under an ESA research, development and test programme. The DS4G concept is able to operate at very high specific impulse and thrust density values well in excess of conventional 3-grid ion thrusters at the expense of a higher power-to-thrust ratio. This makes it a possible candidate for ambitious missions requiring very high delta-V capability and high power. Such missions include 100 kW-level multi-ton probes based on nuclear and solar electric propulsion (SEP) to distant Kuiper Belt Object and inner Oort cloud objects, and to the Local Interstellar medium. In this paper, the DS4G concept is introduced and its application to this mission class is investigated. Benefits of using the DS4G over conventional thrusters include reduced transfer time and increased payload mass, if suitably advanced lightweight power system technologies are developed.A mission-level optimisation is performed (launch, spacecraft system design and low-thrust trajectory combined) in order to find design solutions with minimum transfer time, maximum scientific payload mass, and to explore the influence of power system specific mass. It is found that the DS4G enables an 8-ton spacecraft with a payload mass of 400 kg, equipped with a 65 kW nuclear reactor with specific mass 25 kg/kW (e.g. Topaz-type with Brayton cycle conversion) to reach 200 AU in 23 years after an Earth escape launch by Ariane 5. In this scenario, the optimum specific impulse for the mission is over 10,000 s, which is well within the capabilities of a single 65 kW DS4G thruster. It is also found that an interstellar probe mission to 200 AU could be accomplished in 25 years using a “medium-term” SEP system with a lightweight 155 kW solar array (2 kg/kW specific mass) and thruster PPU (3.7 kg/kW) and an Earth escape launch on Ariane 5. In this case, the optimum specific impulse is lower at 3500 s which is well within conventional gridded ion thruster capability.     

Satellite orbit perturbations in a dusty Martian atmosphere

In this paper we calculate the effect of atmospheric dust on the orbital elements of a satellite. Dust storms that originate in the Martian surface may evolve into global storms in the atmosphere that can last for months can affect low orbiter and lander missions. We model the dust as a velocity-square depended drag force acting on a satellite and we derive an appropriate disturbing function that accounts for the effect of dust on the orbit, using a Lagrangean formulation. A first-order perturbation solution of Lagrange's planetary equations of motion indicates that for a local dust storm cloud that has a possible density of 8.323×10⁻¹⁰ kg m⁻³ at an altitude of 100 km affects the orbital semimajor axis of a 1000 kg satellite up −0.142 m day⁻¹. Regional dust storms of the same density may affect the semimajor axis up to of −0.418 m day⁻¹. Other orbital elements are also affected but to a lesser extent.