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.     

Technology-based design and scaling for RTGs for space exploration in the 100 W range

This paper presents the results of a study on design considerations for a 100 W radioisotope thermo-electric generator (RTG). Special emphasis has been put on designing a modular, multi-purpose system with high overall TRL levels and making full use of the extensive Russian heritage in the design of radioisotope power systems. The modular approach allowed insight into the scaling of such RTGs covering the electric power range from 50 to 200 W_e (EoL). The retained concept is based on a modular thermal block structure, a radiative inner-RTG heat transfer and using a two-stage thermo-electric conversion system.     

Aerobraking at Venus: A science and technology enabler

Venus remains one of the great unexplored planets in our solar system, with key questions remaining on the evolution of its atmosphere and climate, its volatile cycles, and the thermal and magmatic evolution of its surface. One potential approach toward answering these questions is to fly a reconnaissance mission that uses a multi-mode radar in a near-circular, low-altitude orbit of ∼400 km and 60–70° inclination. This type of mission profile results in a total mission delta-V of ∼4.4 km/s. Aerobraking could provide a significant portion, potentially up to half, of this energy transfer, thereby permitting more mass to be allocated to the spacecraft and science payload or facilitating the use of smaller, cheaper launch vehicles.Aerobraking at Venus also provides additional science benefits through the measurement of upper atmospheric density (recovered from accelerometer data) and temperature values, especially near the terminator where temperature changes are abrupt and constant pressure levels drop dramatically in altitude from day to night.Scientifically rich, Venus is also an ideal location for implementing aerobraking techniques. Its thick lower atmosphere and slow planet rotation result in relatively more predictable atmospheric densities than Mars. The upper atmosphere (aerobraking altitudes) of Venus has a density variation of 8% compared to Mars' 30% variability. In general, most aerobraking missions try to minimize the duration of the aerobraking phase to keep costs down. These short phases have limited margin to account for contingencies. It is the stable and predictive nature of Venus' atmosphere that provides safer aerobraking opportunities.The nature of aerobraking at Venus provides ideal opportunities to demonstrate aerobraking enhancements and techniques yet to be used at Mars, such as flying a temperature corridor (versus a heat-rate corridor) and using a thermal-response surface algorithm and autonomous aerobraking, shifting many daily ground activities to onboard the spacecraft. A defined aerobraking temperature corridor, based on spacecraft component maximum temperatures, can be employed on a spacecraft specifically designed for aerobraking, and will predict subsequent aerobraking orbits and prescribe apoapsis propulsive maneuvers to maintain the spacecraft within its specified temperature limits. A spacecraft specifically designed for aerobraking in the Venus environment can provide a cost-effective platform for achieving these expanded science and technology goals.This paper discusses the scientific merits of a low-altitude, near-circular orbit at Venus, highlights the differences in aerobraking at Venus versus Mars, and presents design data using a flight system specifically designed for an aerobraking mission at Venus. Using aerobraking to achieve a low altitude orbit at Venus may pave the way for various technology demonstrations, such as autonomous aerobraking techniques and/or new science measurements like a multi-mode, synthetic aperture radar capable of altimetry and radiometry with performance that is significantly more capable than Magellan.     

Endurance testing of a pulsed plasma thruster for nanosatellites

The mission complexity of Nanosatellites has increased tremendously in recent years, but their mission range is limited due to the lack of an active orbit control or ∆v capability. Pulsed Plasma Thrusters (PPT), featuring structural simplicity and very low power consumption are a prime candidate for such applications. However, the required miniaturization of standard PPTs and the adaption to the low power consumption is not straightforward. Most investigated systems have failed to show the required lifetime. The present coaxial design has shown a lifetime of up to 1 million discharges at discharge energies of 1.8 J in previous studies. The present paper focuses on performance characterizations of this design. For this purpose direct thrust measurements with a µN thrust balance were conducted. Thrust measurements in conjunction with mass bit determination allowed a comprehensive assessment. Based on those measurements the present µPPT has a total impulses capability of approximately I≈1.7 Ns, an average mass bit of 0.37 µg s⁻¹ and an average specific impulse of I_sp≈904 s. All tests have shown very good EM compatibility of the PPT with the electronics of the flight-like printed circuit board. Consequently, a complete µPPT unit can provide a ∆v change of 5.1 m/s or 2.6 m/s to a standard 1-unit or 2-unit CubeSat respectively.     

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.     

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.     

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.     

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.     

MOA2—an R&D paradigm buster enabling space propulsion by commercial applications

More than 60 years after the late Nobel laureate Hannes Alfvén had published a letter stating that oscillating magnetic fields can accelerate ionised matter via magneto–hydrodynamic interactions in a wave like fashion, the technical implementation of Alfvén waves for propulsive purposes has been proposed, patented and examined for the first time by a group of inventors.Consequently improved since then, the name of the latest concept, relying on magneto-acoustic waves to accelerate electric conductive matter, is MOA²—Magnetic field Oscillating Amplified Accelerator. Based on computer simulations, which were undertaken to get a first estimate on the performance of the system, MOA² is a corrosion free and highly flexible propulsion system, whose performance parameters might easily be adapted in operation, by changing the mass flow and/or the power level. As such the system is capable of delivering a maximum specific impulse of 13116 s (12.87 mN) at a power level of 11.16 kW, using Xe as propellant, but can also be attuned to provide a thrust of 236.5 mN (2411 s) at 6.15 kW of power. First tests—that are further described in this paper—have been conducted successfully with a 400 W prototype system at an ambient pressure of 0.20 Pa, delivered 9.24 mN of thrust at 1472 s I_SP, thereby underlining the feasibility of the concept.Based on these results, space propulsion is expected to be a prime application for MOA²—a claim that is supported by numerous applications such as Solar and/or Nuclear Electric Propulsion or even as an ‘afterburner system’ for Nuclear Thermal Propulsion. However, MOA² has so far seen most of its R&D impetus from terrestrial applications, like coating, semiconductor implantation and manufacturing as well as steel cutting. Based on this observation, MOA² resembles an R&D paradigm buster, as it is the first space propulsion system, whose R&D is driven primarily by its terrestrial applications. Different terrestrial applications exist, but the most successful scenarios so far revolve around MOA²'s unique features with respect to high throughput/low target temperature coatings on sensitive materials. In combination with its intrinsic high flexibility, MOA² is highly suited for a common space-terrestrial application research and utilisation strategy.This paper presents the recent developments of the MOA² R&D activities at Q₂ Technologie(s), the company in Vienna, Austria, which has been set up to further develop and test the magneto-acoustic wave technology and its applications.     

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.     

Comparison of rainfall models with Ku-band beacon measurement

Predictions of rain rate and rain attenuation are the most vital steps when analyzing a satellite link operating at frequencies above 10 GHz. Rain attenuation at 12.594 GHz over a satellite path link was measured for the period of 3 years (i.e. January 2002 to December 2004) at Bangkok (13.7°N, 100.7°E). In this paper, a comparison between the current methodologies available to model the impact of rain in earth-space propagation and a dataset of 3 years of rain accumulation with a sampling period of 1 min is made.     

Hypothetical flora and fauna of Venus

Hypothetical habitability of some of extrasolar planets is a fundamental question of science. Some of exoplanets possess physical conditions close to those of Venus. Therefore, the planet Venus, with its dense and hot (735 K) oxygen-free atmosphere of CO₂, having a high pressure of 9.2 MPa at the surface, can be a natural laboratory for this kind of studies. The only existing data on the planet׳s surface are still the results obtained by the Soviet VENERA landers in the 1970s and 1980s. The TV experiments of Venera-9 and 10 (October, 1975) and Venera-13 and 14 (March, 1982) delivered 41 panoramas of Venus surface (or their fragments). There have not been any similar missions to Venus in the subsequent 39 and 32 years. In the absence of new landing missions to Venus, the VENERA panoramas have been re-processed. The results of these missions are studied anew. A dozen of relatively large objects, from a decimeter to half a meter in size, with an unusual morphology have been found which moved very slowly or changed slightly their shape. Their emergence by chance could hardly be explained by noise. Certain unusual findings that have similar structure were found in different areas of the planet. This paper presents the last results obtained of a search for hypothetical flora and fauna of Venus.     

Design and testing of a roto-translational shutter mechanism for planetary operation

This work describes the design and testing of a shutter mechanism for a miniaturized infrared spectrometer developed for the ESA ExoMars Pasteur mission. Unlike most usual cover mechanisms, the conceived one provides a roto-translational motion. This feature allows the sealing of the interferometer main entrance window from dust contamination, in addition to the usual function of shuttering the instrument field of view. Although this characteristic is strongly desired because it avoids dust deposition and optics contamination while the instrument is not operating, it makes the mechanism design significantly more complex. Moreover, challenging design constraints were faced: the mass budget allowed for no more than 30 g allocation, the expected working thermal range extended down to −80 °C and high vibration levels with an acceleration peak of 670 m/s² were predicted during Mars landing. To complete the picture, the mechanism cover was required to provide also a calibration target for the 2–25 μm spectral range of the spectrometer. The resulting system is made by a calibrating/shutter cover moved by a purposely designed out of plane cams system which provides the desired motion. A mechanism mockup was assembled and successfully tested in the predicted thermal and mechanical environments.     

Fuzzy logic active flatness control of a space membrane structure

This paper presents an experimental study on wrinkle reduction in a rectangular membrane structure with elliptical boundary cuts. The structure is subjected to two different heat loads: 145 °C and 205 °C. Wrinkling induced by the thermal expansion of the membrane is measured using a photogrammetry system and the root mean square of the membrane's surface above a plane of best fit. A fuzzy logic controller is developed to track a desired flatness level when the membrane is subject to either application or removal of a thermal load. Experimental verification is conducted using a membrane structure test facility and the results show that wrinkles have been reduced significantly and quickly using the proposed fuzzy logic controller.     

Investigation of nuclear electric powered interstellar precursor missions

Nuclear Electric Propulsion (NEP) is a technology conceptually proposed since the 1940s by E. Stuhlinger in Germany. The JIMO mission originally planned by NASA in the early 2000s produced at least two designs of ion thrusters fed by a 20–30 kW nuclear powerplant. When compared to conventional (chemical) propulsion, the major advantage of NEP in the JIMO context was recognized to be the much higher I_sp (lab-tested at up to 15,000 s) and the capability for sustained power generation, up to 8–10 years when derated to I_sp about 8000 s.The goal of this paper is to show that current or near term NEP technology enables missions far beyond our immediate interplanetary backyard. In fact, by extending the semi-analytical approach used by Stuhlinger, with reasonable ratios α≡power/mass of the propulsion system (i.e., 0.1– 0.4 kW/kg), missions to the Kuiper Belt (40 AU and beyond) and even the so-called FOCAL mission (at 540 AU) become feasible with an attractive payload fraction and in times of order 10–15 years.Further results regarding missions to Sedna’s perihelion/aphelion, and to Oort’s cloud will also be presented, showing the constraints affecting their feasibility and mass budget.     

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.     

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.     

MESSENGER at Mercury: Early orbital operations

The MErcury Surface, Space ENvironment, GEochemistry, and Ranging (MESSENGER) spacecraft, launched in August 2004 under NASA's Discovery Program, was inserted into orbit about the planet Mercury in March 2011. MESSENGER's three flybys of Mercury in 2008–2009 marked the first spacecraft visits to the innermost planet since the Mariner 10 flybys in 1974–1975. The unprecedented orbital operations are yielding new insights into the nature and evolution of Mercury. The scientific questions that frame the MESSENGER mission led to the mission measurement objectives to be achieved by the seven payload instruments and the radio science experiment. Interweaving the full set of required orbital observations in a manner that maximizes the opportunity to satisfy all mission objectives and yet meet stringent spacecraft pointing and thermal constraints was a complex optimization problem that was solved with a software tool that simulates science observations and tracks progress toward meeting each objective. The final orbital observation plan, the outcome of that optimization process, meets all mission objectives. MESSENGER's Mercury Dual Imaging System is acquiring a global monochromatic image mosaic at better than 90% coverage and at least 250 m average resolution, a global color image mosaic at better than 90% coverage and at least 1 km average resolution, and global stereo imaging at better than 80% coverage and at least 250 m average resolution. Higher-resolution images are also being acquired of targeted areas. The elemental remote sensing instruments, including the Gamma-Ray and Neutron Spectrometer and the X-Ray Spectrometer, are being operated nearly continuously and will establish the average surface abundances of most major elements. The Visible and Infrared Spectrograph channel of MESSENGER's Mercury Atmospheric and Surface Composition Spectrometer is acquiring a global map of spectral reflectance from 300 to 1450 nm wavelength at a range of incidence and emission angles. Targeted areas have been selected for spectral coverage into the ultraviolet with the Ultraviolet and Visible Spectrometer (UVVS). MESSENGER's Mercury Laser Altimeter is acquiring topographic profiles when the slant range to Mercury's surface is less than 1800 km, encompassing latitudes from 20°S to the north pole. Topography over the remainder of the southern hemisphere will be derived from stereo imaging, radio occultations, and limb profiles. MESSENGER's radio science experiment is determining Mercury's gravity field from Doppler signals acquired during frequent downlinks. MESSENGER's Magnetometer is measuring the vector magnetic field both within Mercury's magnetosphere and in Mercury's solar wind environment at an instrument sampling rate of up to 20 samples/s. The UVVS is determining the three-dimensional, time-dependent distribution of Mercury's exospheric neutral and ionic species via their emission lines. During each spacecraft orbit, the Energetic Particle Spectrometer measures energetic electrons and ions, and the Fast Imaging Plasma Spectrometer measures the energies and mass per charge of thermal plasma components, both within Mercury's magnetosphere and in Mercury's solar-wind environment. The primary mission observation sequence will continue for one Earth year, until March 2012. An extended mission, currently under discussion with NASA, would add a second year of orbital observations targeting a set of focused follow-on questions that build on observations to date and take advantage of the more active Sun expected during 2012–2013. MESSENGER's total primary mission cost, projected at $446 M in real-year dollars, is comparable to that of Mariner 10 after adjustment for inflation.     

SLUCUBE: Innovative high performance nanosatellite science and technology demonstration mission

The Space Systems Research Laboratory (SSRL) at Saint Louis University is developing SLUCUBE nanosatellite as part of the space mission design program. The objective of the mission is to demonstrate space capability of high performance nanosatellite components that has been developed at SSRL for the past three years. The objective of the program is to provide extremely low-cost and rapid access to space for scientists and commercial exploitation using commercial-off-the-shelf components. SLUCUBE is a double CubeSat with dimensions 10×10×20 cm and a mass of 2 kg. This nanosatellite features suite of technology demonstration components to enlarge the capability of space mission for such class of spacecrafts. The primary mission of SLUCUBE is to test and demonstrate several enabling technologies by flying a number of university developed high performance components. This paper describes the new developed technologies by providing details of specific components developed along with the R&D efforts and laboratory facilities. A brief discussion about the student involvement and educational benefits will also be presented.