Power Instruments & Planetary Infrastructure

To enable any development to reach as wide a market as possible



The FSP systems could also provide power for ISRU plants and habitats  benefits



SPACE POWER APPLICATIONS

Space fission nuclear power can benefit a wide range of applications and, in specific cases, can be the mission enabling technology

Two sizes of generator considered: 30 kWe and 200 kWe

The key issue is the rationale for identifying common features to enable any development to reach as wide a market as possible



The key issues are the rationale for using

fission nuclear power and identifying common

features to enable any development to reach as

wide a market as possible 


1.1       Introduction


  Some applications can in principle be appropriate to both sizes but not necessarily with the same technology.


Applications are considered in terms of:

•          High power instruments such as radars or lasers,

•          Planetary infrastructure.




1.2       High Power Instruments (radar, laser, etc.)


1.2.1    Background

Advances in solar power and electronics allow high power radars in earth orbit without the need for nuclear generators.  Performance has also been enhanced by better signal processing techniques.  Exploration of distant planets however can be greatly enhanced by the use of very high power ground penetrating radars which would need a high power generator.  The Russian RORSAT programme gave a basic capability with 3 kW of power and this would be a useful level; but higher power, up to say 30 kW, would enable investigation of the sub-surface structures.


A radar sounder’s performance depends on its ability to penetrate deep in the subsurface. Deep penetration is especially critical for the detection of water/ice interface in Europa whose depth is uncertain. The ability to penetrate in the sub-surface is directly dependent on the radiated power by the radar. Dynamic range of MARSIS and SHARAD - two radars currently operating at Mars - is limited to ~ 50 dB dictated by the galactic background noise and radar’s radiated power.  One method to increase the dynamic range further is to increase the radiated power. Both MARSIS and SHARAD have a radiated power in the order of 10 We.  An increase of this radiated power to 3-30 kW (achievable with nuclear power generation) will increase the dynamic range by 25-35 dB (Safaeinili et al. 2005), allowing deeper signal penetrations.


In most orbiting radar, the transmitted waveform is a chirp, a long pulse that is linearly modulated in frequency. Because the bandwidth of the chirp can be at most equal to the highest frequency of the radar pulse (although for practical purposes this limit is never reached), it can be seen that vertical resolution requires the radar to operate at high frequency.  Thus, penetration and resolution are conflicting requirements, and no orbiting subsurface sounding radar can achieve one without compromising on the other.


1.2.2   MARSIS and SHARAD

An example of this tradeoff is the design of the two above mentioned radars currently operating at Mars: MARSIS and SHARAD. They are both synthetic-aperture, orbital sounding radars, carried respectively by ESA’s Mars Express and NASA’s Mars Reconnaissance Orbiter.  MARSIS is capable of transmitting at four different bands between 1.3 MHz and 5.5 MHz, with a 1 MHz bandwidth. SHARAD operates at a central frequency of 20 MHz transmitting a 10 MHz bandwidth.


Whereas MARSIS is optimized for deep penetration, having detected echoes down to a depth of 3.7 km over the South Polar Layered Deposits (Plaut et al. 2007), SHARAD is capable of a tenfold-finer vertical resolution, namely 15 m or less, depending on the dielectric constant of the material being sounded.  The difference in the performance of the two radars can be appreciated through the examination of Figs. below. Both figures show radargrams that are representations of radar echoes acquired continuously during the movement of the spacecraft as grey-scale images, in which the horizontal dimension is distance along the ground track, the vertical dimension is the round trip time of the echo (Figure 2) or depth (Figure 3), and the brightness of the pixel is a function of the strength of the echo.


The capability of MARSIS for deep penetration at the expense of vertical resolution can be seen in the radargram in Figure 3, in which subsurface echoes almost as bright as surface ones are detected down to depths of 3.5 km, but only very faint details of the stratigraphy can be discerned.

In Figure 3 the tenfold-greater vertical resolution of SHARAD, achieved at the expense of penetration, is made evident by the richness of details discernible in the subsurface stratigraphy.  Remarkably, it should be noted that a serious reason of concern for the performance of a radar sounder for the icy moons of the Solar System is the noise introduced by the giant planets’ radio emission at decametric and hectometric wavelengths, especially Jupiter’s (see e.g. Zarka et al. 2004). However, the radio noise from Jupiter should not impact the radar performance on the anti-Jovian side of a moon.


Just for comparison, if the peak power during pulse transmission is reduced to 10W (typical value for MARSIS and SHARAD radar embedded on ESA’s Mars Express and NASA’s Mars Reconnaissance Orbiter) and all the other parameters are kept fixed, the acceptable losses inside ice are reduced to 75 dB, limiting penetration to 4.5 to 19 km, depending on the impurity content of Europa ice.


In Table 1 (Moore 2000) the attenuation for different types of impurities in ice is shown, based on laboratory measurements, ice temperature modelling for Europa and some scaling from Earth ice measurements. These data are valid for electromagnetic frequencies of a few tens of MHz.

In this table, the first column (M) represents the plausibility of the ice type for Europa (second column) given the present understanding: 0 is the least likely while 3 is the more likely; the third column indicates the impurity content, the major responsible for radar wave attenuation; the fourth column is the attenuation in dB/m at 251 K; the columns I, II and III are computed two-way attenuations, in dB/km for ice shells with base temperatures of 270, 260 and 250 K. The range of values under these columns corresponds to surface temperatures of 50 and 100 K, respectively.


Concluding, the adaption of a 30 kWe nuclear reactor power generator for a mission towards Jovian icy moon Europa could, among other advantages, enhance the penetration capabilities of a sounding radar with respect to those at the moment obtainable from current power generators for this kind of missions (RPSs) from a minimum of 2.5 km to a maximum of 8 km, under the assumptions above mentioned.


1.3       Planetary Infrastructure (lunar or Mars base)


1.3.1    Fission Surface Power (FSP)


Reliability and countering sunlight limitations is why Nuclear Power is uniquely the best system to provide high power to surface space missions requiring in difficult environments. Operational robustness would be achieved through multi-fault tolerant design architectures and long-life components, requiring no maintenance. These are the conditions that are anticipated for future human missions to the Moon and Mars.


A key advantage of FSP systems is that they produce constant power to allow continuous day and night surface operations. In fact, nuclear systems are likely to produce higher power during the night, due to the colder thermal sink, when heating and lighting requirements are greatest. Relative to comparable solar power systems with energy storage, FSP offers significant mass and volume savings. Based on recent NASA studies, the cost of nuclear power systems is also competitive with these solar-based power systems.


The moon’s 29.5 days’ rotational period results in a long, cold, lunar night of 354 hours. On Mars, the night period is only 12 hours but sunlight at the surface is reduced to about 20 % that of the Moon. Martian dust storms and missions to the higher latitudes, further decrease the availability of sunlight for solar power. On either surface, crew members would be highly dependent on the power system to achieve mission objectives and assure human safety.


So, Nuclear system designs can be highly flexible and operationally robust. By burying the reactor in the local terrain, the system could be located close to the habitation area with minimal radiation concerns. If desired, the system could include on-board shielding for above-grade installation and high-voltage cabling to permit connection to (or separation from) remote infrastructure.

1.3.2   In situ resource utilization (ISRU)


The FSP systems could also provide power for ISRU plants and habitats with the following benefits:


 1.3.2.1           ISRU – BENEFITS


The following have been identified as potential ISRU benefits:

•          Mass Reduction (delivered mass to surface)

o          In-situ production of mission critical consumables (propellants, life

            support consumable, and fuel cell reactants)

o          Habitat shielding (radiation, micrometeoroid&exhaust plumedebris)

            and surface nuclear power (radiation) from in-situ materials (raw or

            processed)

o          For sustained human presence manufacturing and construction of

            infrastructure

•          Cost Reduction (reduction in mission costs)

o          Reduction of mass and reuse of elements

o          Use of modular, common hardware with propulsion, life support & fuel cell power systems

o          ISRU enabled missions lead to reduction in architecture costs through elimination of separate dedicated missions

o          Cost reduction through commercial sector participation

o          Risk Reduction & Mission Flexibility

•          Reduction in mission risk due to:

o          Reduction in Earth launches and sequential mission events

o          Surface manufacturing & repair

o          Dissimilar redundancy of critical system

o          Reduction in mission & crew risk due to increased shielding

o          High mission flexibility - use of common modular h/w-consumables

•          Mission Enhancements and Enabled Capabilities

o          Increased robotic and human surface access through hoppers

o          Increased delivered and return payload mass through ISRU

o          Lower cost Moon missions with in-space depots and lunar delivered

            propellant

o          Energy-rich and extended missions through production of mission

            consumables and power

o          Low-cost mass-efficient manufacturing, repair, and habitation & power infrastructure growth

Studies anticipated that a small manned lunar or Martian base could be powered by a nuclear generator of the order of 20 – 30 kWe.  It would also make good use of the waste heat although probably not all of the 500-600 kWth that might be expected.


Arguably a generator of this size could also power an un- (or periodically) inhabited planetary base or possibly a space port at the second sun earth Lagrange point for assembling and maintaining large space structures.  Power requirements for human-tended surface outposts and bases are expected to range from 25 to 100 kWe during the early build-up phases. As the base becomes fully operational with in-situ resource production and closed-loop life support, power requirements could approach 1 MWe.

Figure 5 shows a representative power requirements profile for the early phases of a potential lunar surface mission that result in a total power requirement of about 70 kWe within two years of establishing the outpost.

1.3.3   Space structures at Lagrangian points


Lagrangian points, or libration points, for two celestial bodies in mutual revolution are the five points (L1 through L5) where a smaller body (third body) can orbit at a fixed distance from the larger masses. These points mark positions where gravitational attraction of the two large masses equals the centripetal force required to rotate with them. Each pair of masses – Earth and Moon, Earth and Sun – has its own set of Lagrange points. The three points (L1, L2, L3) lying along a line joining the two larger masses are positions of unstable equilibrium; L1 and L2 points are unstable (on a time scale of approximately 60 days for Sun-Earth system), which requires satellites and in prospective space-ports parked at these positions to undergo regular station keeping. L4 and L5 points are positions of stable equilibrium, as long as mass ratio between the large masses exceeds 24.96. This condition is satisfied for both Earth-Sun, Earth-Moon systems and many others pairs of bodies in the solar system. Figure 6 shows Sun-Earth Lagrangian points; halo orbits, periodic orbits around SEL1 and SEL2 points are also depicted.

The L1 point of the Sun-Earth system affords an uninterrupted view of the Sun and currently SOHO, ACE, WIND satellites study Sun from halo orbit around this point. The SEL2 point is a good point for space-base observatories; currently NASA’s WMAP spacecraft, ESA’s Herschel and Planck Space observatories, CNSA’s Change’s 2 are in orbit around this point.


Whereas several space missions have been launched to the two collinear equilibrium points L1 and L2, taking advantage of their dynamical and geometrical characteristics, the region around L3 is so far unexploited. This is essentially due to the severe communication limitations caused by the distant and permanent opposition to the Earth, and by the gravitational perturbations mainly induced by Jupiter and the close passages of Venus, whose effects are more important.


The L4 and L5 points are home to stable orbits; this causes dust and space debris accumulation. A natural example of this phenomenon are the 2,500 Trojan asteroids ensnared at Jupiter-Sun’s L4 and L5 points; large concentrations of dust at Moon-Earth L4 and L5 points were discovered by Kordylewsky in 1956.

1.3.4   On-orbit assembly and servicing facilities


Several studies have examined facilities near the Earth-Moon and Sun-Earth Lagrange points.  A great deal of work has been done regarding the on-orbit assembly of large aperture telescopes. Villard (2004) have proposed assembling a Very Large Space Telescopes at a facility located at the first Earth-Moon Lagrange point (L1), while Oegerle et al. (2006) proposed assembly of a large scalable space observatory (10-50 meter segmented filled-aperture) at the Earth-Moon L2 point to be afterwards transported to the Sun-Earth L2 point.


A facility at the Earth-Moon L1/L2 point (~60,000 km from the Moon on the Earth-Moon line) would be a convenient staging point for travel to the Moon, Mars and Sun-Earth Lagrangian points for very little Delta-V. In addition to providing a mission staging and crew habitation platform for expeditions to the lunar surface, the gateway could support travel to many solar system destinations, including Mars, and the Sun-Earth Lagrange points. It could also serve as a test bed for developing technologies, systems, and operations procedures for future exploration missions, and a facility for the assembly, checkout and maintenance of future space observatories.


Figure 7 shows one conceptual design for such a facility. It consists of a rigid core ~4-m in diameter and 11-m long, with docking ports; air locks for access; propulsion system for manoeuvring; avionics electrical power for attitude control, thermal control, command and data handling; life support systems; an EVA work platform for telescope assembly; EVA toolbox; and a remote manipulator system. An inflatable torus with a 3-meter diameter cross-section wrapped around the core provides a pressurized volume of ~135 m3 for crew quarters. When completed, the facility would have a total volume of ~275m3and a mass of ~23,400 kg. The facility is designed to accommodate a crew of 4-6 people for two 10 to 30 day missions each year, with a 15 years’ design lifetime. Such a facility could be powered by a nuclear fission generator only as estimated power budget is 12 kWe.


If such a facility would be developed for the exploration program, future space observatories could use a few 10’s of m/s of Delta-V of its propulsion budget to travel from an Sun-Earth L2 orbit (a good point for space-based observatories) to the Earth-Moon L1 facility for servicing. A space tug would likely be used to manoeuvres space observatory within reach of the servicing arm at the gateway, and later to send it back toward Sun-Earth L2. Employing this approach estimates indicate that servicing an SAFIR-like telescope (Single-Aperture Far-Infrared Observatory) could extend operational lifetime several-fold and 1 or 2 orders-of-magnitude improvement in its performance could be achieved (Lillie, 2006).

1.3.5   US Studies


After NASA’s cancellation of the Jupiter Icy Moons Orbiter (JIMO) mission, several study groups explored the application of fission power systems for lunar and Mars surface missions.  One of them, under the direction of the Exploration Systems Architecture Study (ESAS), developed a lunar fission surface power system derived from the JIMO concept. It used a gas-cooled reactor and direct Brayton power conversion at 1150 K turbine inlet temperature. Like JIMO, this concept required advanced reactor fuel and refractory cladding materials due to the high operating temperatures.


Another group, under the NASA Exploration Technology Development Program (ETDP) and in partnership with the Department of Energy (DOE), conducted a 60-day study that recommended a low temperature, liquid-metal cooled reactor using more conventional UO2 fuel and stainless steel cladding. The low temperature reactor was proposed in order to reduce development risk and cost. The 60-day study team also recommended continued evaluation of three different power conversion technologies: Brayton, Stirling, and thermoelectric.

Further studies presented comparisons of reactor and power conversion design options for 50 kWe class lunar and Mars surface power applications with scaling from 25 to 200 kWe. These studies examined the mass and performance of low temperature, stainless steel based reactors and higher temperature refractory reactors. The preferred system implementation approach used crew-assisted assembly and in-situ radiation shielding via installation of the reactor in an excavated hole. As an alternative, self-deployable system concept that use earth-delivered, on-board radiation shielding was evaluated. The analyses indicated that among the 50 kWe stainless steel reactor options, the liquid-metal Stirling system provides the lowest mass at about 5300 kg followed by the gas-cooled Brayton at 5700 kg and the liquid-metal thermoelectric at 8400 kg. The use of a higher temperature, refractory reactor favours the gas-cooled Brayton option with a system mass of about 4200 kg as compared to the Stirling and thermoelectric options at 4700 and 5600 kg, respectively. The self-deployed concepts with on-board shielding resulted in a factor of two system mass increase as compared to the in-situ shielded concepts.


It was also recognized that lunar mission presents a more difficult thermal environment for waste heat radiators, given the higher solar isolation levels and longer daylight periods, while the carbon-dioxide atmosphere at Mars introduces potential material issues, especially for refractory based reactor concepts. However, this might be managed through the use of a containment structure surrounding the exposed refractory components and an inert cover gas.


Three leading reactor options for space fission power applications such as the lunar and Mars surface mission – liquid-metal cooled, gas cooled, and heat pipe cooled – are at the moment envisaged. The reactor could be fast-spectrum or moderated and the fuel and core construction material would dictate operating temperature. Possible fuel options for the low temperature reactor included UO2 and UZrH. If refractory alloys, such as niobium, tantalum, or molybdenum, could be used for the fuel cladding and primary coolant boundary, operating temperatures up to about 1300 K could be considered. At the higher temperature, UO2 and UN were assumed as the likely fuel choices. The refractory reactor introduces additional risk given the greater uncertainty for long-term material performance, but also the potential for greater power conversion efficiency, smaller waste heat radiators, and reduced system mass.

Two fundamental technological issues were also taken into account: reactor shielding and heat rejection.  For what concerns the first issue, two different options were analyzed: an excavation shield and an on-board, Earth-delivered shield. The preferred approach was to place the reactor, enclosed by an upper conical shadow shield, comprised of Wolframium and Lithium Hydride, and a 2 cm thick boron aluminium liner in an excavated hole to provide crew-rated radiation protection (radiation dose to less than 5 rem/year at a radial distance of 6 meters from the reactor center line).


The second alternative consisted of a circumferential side shield and an inverted cone top shield, both assembled with alternating layers of Wolframium and Lithium Hydride. The reactor was placed directly on the regolith surface and a Boral bottom plate limited neutron back-scatter. To reduce mass, the circumferential side shield was sectored with a 90° segment sized to limit radiation to 5 rem/yr at 2 km in the direction of the crew habitat area. The remaining 270° portion, as well as the top shield, is sized for 50 rem/yr at 2 km, permitting safe short term operations by crew members and acceptable radiation levels for the power conversion equipment. The 2 km separation distance was determined based on an optimization of shield mass versus power cabling mass.


The other issue also taken into account was heat rejection, in fact waste heat from the power converters must be removed and transported to radiator panels where it can be rejected to the space environment. A pumped-loop heat transport system coupled to a heat pipe radiator, similar to the JIMO concept (Siamidis, et al., 2005), was proposed. The heat pipe radiator approach permits efficient heat spreading to the panel surface and reduces the vulnerability to micrometeoroid damage as compared to an all pumped-loop system.


The surface thermal environment has a significant effect on radiator performance, especially during daylight conditions. Horizontal radiators with one-sided heat rejection view deep space and absorb direct solar radiation during the day. Using high emissivity, low absorbtivity coatings would result in equivalent radiator sink temperatures of about 262 K at solar noon. Vertical radiators would absorb both direct solar radiation and reflected surface radiation resulting in an equivalent sink temperature of 317 K at solar noon. However, the vertical radiator would have a smaller platform area given its ability to reject heat from two sides.

1.3.6   European STUDIES


Europe has neither developed nor used fission reactors in space so far. However, based on important expertise from terrestrial and naval reactors, substantial design work for space reactors has been performed since the ‘70s, especially in France, Germany, and the UK and within ESA. Most of the recent efforts focused on lunar surface reactors of smaller dimensions (5–100 kWe) and less ambitious technical designs than either the US SP-100 (100 kWe) or Prometheus (200 kWe) reactors.


Instead of a technically optimal system requiring substantial nuclear and material developments, the approach was to accept some mass penalties, but base the designs on standard nuclear fuel and, where possible, terrestrial technologies for all other key system components, except the cold well.


The driving idea of a recent European study (Finzi et al. 2007) for a lunar surface reactor was to adapt, as much as possible, technology from the well-established terrestrial pressurized water reactor (PWR) technology to the design of a reactor suited for space conditions.  An integral type reactor with a Rankine steam cycle was chosen as reference system. The concept used a novel and innovative reactivity control by changing the core geometry. In order to adapt terrestrial PWR designs to space, the following main changes were introduced: U-ZrH1.7 matrix fuel (45 per cent: 93 per cent enriched uranium and 55 per cent ZrH1.7) (the same fuel was used in the US SNAP 10A reactor (in principle thus space qualified) and well-known material tested in terrestrial TRIGA reactors); 0.3 mm stainless steel cladding; maximum temperature of 345◦C (about 15◦C higher than terrestrial PWRs); and minimum inlet temperature of 335◦C (45◦C higher than terrestrial PWRs). In Figure 11 a sketch of the primary vertical cross section is depicted.

Two geometries were considered for the fuel rods positioning (configuration A: standard hexagonal fuel rods and configuration B: an innovative hexagonal fuel rod with internal cooling holes). In all configurations, the cold well was the main mass driver with 2500 kg of the total 4300 kg mass, whereas the reactor vessel including the core and the primary cooling cycle contributed to about 800 kg.


As mentioned above, the electricity production was delegated to a Rankine steam cycle. Preliminary optimization studies revealed a 12% conversion efficiency with the heat sink temperature equal to 165°C.  In Table 2, main system parameters of the 100 kWe lunar surface reactor designed under constraints above illustrated, are shown.