Mars Direct: Humans to the red planet by 1999

“Mars Direct”, is an approach to the space Exploration Initiative that allows for the rapid initiation of manned Mars exploration, possibly as early as 1999. The approach does not require any on-orbit assembly or refueling or any support from the Space Station or other orbital infrastructure. Furthermore, the Mars Direct plan is not merely a “flags and footprints” one-shot expedition, but puts into place immediately an economical method of Earth-Mars transportation, real surface exploratory mobility, and significant base capabilities that can evolve into a mostly self-sufficient Mars settlement. This paper presents both the initial and evolutionary phases of the Mars Direct plan. In the initial phase, only chemical propulsion is used, sendig 4 persons on conjunction class Mars exploratory missions. Two heavy lift booster launches are required to support each mission. The first launch delivers an unfueled Earth Return Vehicle (ERV) to the martian surface, where it fills itself with methane/oxygen bipropellant manufactured primarily out of indigenous resources. After propellant production is completed, a second launch delivers the crew to the prepared site, where they conduct regional exploration for 1.5 years and then return directly to Earth in the ERV. In the second phase of Mars Direct, nuclear thermal propulsion is used to cut crew transit times in half, increase cargo delivery capacity, and to create the potential for true global mobility through the use of CO₂ propelled ballistic hopping vehicles (“NIMFs”). In this paper we present both phases of the Mars Direct plan, including mission architecture, vehicle designs, and exploratory strategy leading to the establishment of a 48 person permanent Mars base. Some speculative thoughts on the possibility of actually colonizing Mars are also presented.     

Parametric study of the diversion of a near Earth object on an Earth intersecting trajectory

To date, NASA's “Near Earth Object Program” has discovered over 5500 comets and asteroids on trajectories that bring them within “the neighborhood” of Earth's orbit. Nearly 1000 of these objects are classified as “potentially hazardous,” passing within 0.05 astronomical units of Earth's orbit. Discovery rates of such threatening bodies increase each year. Given this multitude of threats, in addition to evidence that the planet has absorbed many impacts over its history, it is reasonable to assume that another object will strike the Earth at some point in the future. Consequently, researchers have studied and proposed several mitigation techniques for such an occurrence. This study seeks to determine how effectively the attachment of a tether and ballast mass would divert the trajectory of such threatening objects. Specifically, the study analyzes the effects over time of such a system on objects of varying orbital semimajor axis and eccentricity, using various tether lengths and ballast masses. It was determined that the technique is most effective for NEOs with high eccentricity and small semimajor axis, and that system performance increases as tether length and ballast mass increase.     

Assessment on the feasibility of future shepherding of asteroid resources

Most plausible futures for space exploration and exploitation require a large mass in Earth orbit. Delivering this mass requires overcoming the Earth's natural gravity well, which imposes a distinct obstacle to any future space venture. An alternative solution is to search for more accessible resources elsewhere. In particular, this paper examines the possibility of future utilisation of near Earth asteroid resources. The accessibility of asteroid material can be estimated by analysing the volume of Keplerian orbital element space from which Earth can be reached under a certain energy threshold and then by mapping this analysis onto an existing statistical near Earth objects (NEO) model. Earth is reached through orbital transfers defined by a series of impulsive manoeuvres and computed using the patched-conic approximation. The NEO model allows an estimation of the probability of finding an object that could be transferred with a given Δv budget. For the first time, a resource map provides a realistic assessment of the mass of material resources in near Earth space as a function of energy investment. The results show that there is a considerable mass of resources that can be accessed and exploited at relatively low levels of energy. More importantly, asteroid resources can be accessed with an entire spectrum of levels of energy, unlike other more massive bodies such as the Earth or Moon, which require a minimum energy threshold implicit in their gravity well. With this resource map, the total change of velocity required to capture an asteroid, or transfer its resources to Earth, can be estimated as a function of object size. Thus, realistic examples of asteroid resource utilisation can be provided.     

Bistatic radar as a tool for earth investigation using small satellites

Bistatic radar is a facility for the Earth remote sensing, which uses large spatial diversity between its transmitter and receiver. Nomogram method is proposed to determine the radar's parameters. Analysis of the nomograms has shown that modern onboard radio facilities allow to obtain spatial resolution of about 100 m at the wavelength λ = 3 cm for LEO satellite (H = 350 km). Experiments of bistatic radiolocation of the Earth near the radioshadow zone were provided using telecommunication link “MIR” orbital station — GEO satellite at wavelength λ = 32 cm. For the first time in practice of bistatic radiolocation of the Earth from space reflected signal in radioshadow zone was observed.The analysis of experimental results verified the developed radiophysical model with the value of sea water conductivity σ = 7.0 mo/m and absorption coefficient due to atmospheric oxygen χ = 0.0096±0.0024 dB/km.     

The data compression problem for the “GAIA” astrometric satellite of ESA

Space-based astrometry has a great tradition at ESA. The first space-based astrometric satellite in history, “Hipparcos”, was launched by ESA in 1989 and, in spite of orbital problems, was able to accomplish almost all of its tasks until it was finally shut down in 1993. The results of the Hipparcos mission were published by ESA in 1997 in the form of six CD-ROMs: the Hipparcos Catalogue contains 118,218 entries with median astrometric precision of around 1 milliarcsec, and specific results for double and multiple systems. In practice, Hipparcos drew for the first time the three-dimensional “map” of the spherical region of the Galaxy surrounding the Sun and having a radius of roughly 1,000 light years.

Then, in 1995, ESA launched the study of a new astrometric satellite, named “GAIA” and about a hundred times more powerful than Hipparcos, i.e. with median astrometric precision of around 10 microarcsec. This new satellite is intended to measure the parallaxes of over 50 million stars in the Galaxy, at least for the brightest stars, and this would mean to “draw” the three-dimensional map of the whole Galaxy, reaching out even to the Magellanic Clouds, 180,000 light years away.

The team of European scientists and engineers now designing GAIA, however, is facing hard technological difficulties. One of these is the design and coding of radically new and ultra-powerful mathematical algorithms for the on-board compression of the 50-million-stars data that GAIA will send to Earth from its intended geostationary orbit. Preliminary estimates of the raw data rates from the GAIA focal plane, in fact, are of the order of a few Gigabits per second. To reduce the data stream to the envisaged telemetry link of 1 Megabit per second, on-board data compression with a 1 to 1,000 ratio is the target. Clearly, this is far beyond the capabilities of any lossless compression technique (enabling compression ratios of 1 to some tens), and so some “wise” lossy compression mathematical procedure must be adopted.

In this paper a GAIA-adapted lossy data compression technique is presented, based on the Karhunen-Loève Transform (KLT). The essence of this method was already used by NASA for the Galileo mission when the large antenna got stuck and the mission was rescued by re-programming the on-board computer in terms of the KLT. That transform was officially named ICT — “Integer Cosine Transform” — by the NASA-JPL team led by Dr. Kahr-Ming Cheung. But the KLT here described for GAIA will of course differ from the JPL one in many regards, owing to the advances in computer technology.

Finally, estimates are also given about the possibility of using the KLT for onboard data compression in case GAIA is going to be put into orbit around the Lagrangian point L2 of the Earth-Sun system, and, above all, in case the number of stars to be observed is actually raised from 50 millions to one billion, as ESA currently appears to be likely to pursue.     

The pharao project: Towards a space clock using cold cs atoms

For several years, the “BNM-Laboratoire Primaire du Temps et des Fréquences” has worked on a cold atom frequency standard. With a cesium atomic fountain a resonance line width of 700 mHz has been obtained leading to a short-term stability of 2 × 10⁻¹³ τ^−1/2 down to 2 × 10⁻¹⁵ at 10⁴ s. A first evaluation of the fountain accuracy has been performed resulting in an accuracy of 3 × 10⁻¹⁵, three times better than previously achieved with thermal beams frequency standards. In the atomic fountain, gravity limits the interaction time to ˜1 s, hence the resonance line width to ˜0.5 Hz. A factor of 10 reduction in the line width could be obtained in a micro-gravity environment. The “Centre National d'Etudes Spatiales” (the French space agency), the “BNM-Laboratoire Primaire du Temps et des Fréquences”, the “Laboratoire de l'Horloge Atomique” and the “Laboratoire Kastler Brossel” have set up a collaboration to investigate a space frequency standard using cold atoms: the PHARAO project. A microgravity prototype has been constructed and operated first in the reduced gravity of aircraft parabolic flights in May 1997. It is designed as a transportable frequency standard. The PHARAO frequency standard could be a key element in future space missions in fundamental physics such as SORT (solar orbit relativity test), detection of gravitational waves, or for the realization of a global time scale and a new generation of positioning system.     

Electrophotometrical studies of equatorial and tropical arcs on board the Salyut-6 orbital station

The results from the electrophotometric investigation of the equatorial and tropical ionospheric arcs on board the orbital station “Salyut-6”, carried out with Bulgarian photometer “Duga”, intended for measurements of the self-radiation of the Earth's upper atmosphere in the lines λ = 6300 Å, λ = 5577 Å, λ = 4278 Å and λ = 6563 Å, are analyzed. From the obtained results of the intensity of the measured emissions is established by calculation that the cause of these arcs is the plasma drift downwards, which leads to intensification of the dissociative recombination of the ions O₂⁺ and of the radiative recombination of O⁺.     

Lunar oxygen: Economic benefit for transportation systems and in situ production by fluorination

The use of oxygen produced on the Moon—called “MOONLOX”—is considered as a propellant component for a reusable Earth-Moon transportation system consisting of an aeroassisted orbital transfer vehicle and a lunar bus for lunar descent/ascent. Conditions for economic benefit are discussed and the processing concept of a lunar oxygen plant based on fluorination is presented. It is shown that the necessary mass of supply from Earth for MOONLOX-production is an important parameter, which may not be neglected due to its strong influence on the economy. In the ideal case where no supplies from Earth are required a reduction of up to 50% in masses to be launched into low Earth orbit can be obtained for a typical lunar mission with use of MOONLOX compared to a reference scenario with Earth-derived propellant. Mass-saving decreases, however, significantly with increasing supply from Earth until a critical supply-rate is reached—measured in percentage of MOONLOX-mass produced and consumed—beyond which mass-saving and thus economically promising lunar oxygen production is no longer possible. This critical supply-rate depends on the scenario for MOONLOX-utilization and is much larger in the case of in situ use of MOONLOX on the lunar surface, e.g. as ascent propellant for the lunar bus, than in the case of export for complete refuelling of both space vehicles. The latter scenario therefore requires significantly more autonomy for MOONLOX-production. The reduction of masses to be transported into low Earth orbit and corresponding MOONLOX-consumption define for given specific Earth-to-LEO transportation costs an upper limit on MOONLOX-production costs beyond which economic benefit is not possible. Depending on the MOONLOX-utilization strategy this upper limit varies between 3000 and 55000 $/kg for current Earth-to-LEO transportation costs.     

Effect of weightlessness on posture and movement control during a whole body reaching task

Here are reported preliminary results of the “Synergy” experiment performed aboard the Russian orbital station “MIR” in July 1993 (Altaïr Mission). The experiment was carried out before, during, and after the space flight of two astronauts (S1 and S2). The duration of the flight was 21 days for S1 and 6 month for S2. The subjects were tested during preflight, inflight and postflight. The astronaut subjects were fixed on the ground by the feet. They were asked to pick up a box in front of them on the ground. Two velocities of movement and two distances of the target to be reached were tested. The movement of several small markers placed on the body was recorded on video tape.

Results show that the shape of head and hand trajectories in the sagittal plane remains roughly the same during the flight in spite of the modification of mechanical constraints. Trajectory invariance does not result in joint angular displacement invariance. These data indicate that the planning of the movement takes place in terms of head and hand trajectories rather than joint rotations as it was previously suggested for simple arm reaching movement.     

New results for the space radiation environment of MIR space station obtained by Liulin dosimeter-radiometer. Comparison with LET spectrometer NAUSICAA

Since 1988 high sensitivity semiconductor dosimeter-radiometer “Liulin” worked on board of MIR space station. Device measured the absorbed dose rate and the flux of penetrating particles. The analysis of the data hows the following new results:

In October 1989 and after March 24, 1991, two additional stable maximums in flux channel were observed in the southern-eastern part of South Atlantic Anomaly (SAA). These two maximums existed at least several months and seem to be due to trapped high energy electron and proton fluxes. In April 1991 additional maximums were localized in the following geographical coordinates regions: LATITUDE = (−35 °)–(−50 °) LONGITUDE = 332 ° − 16 ° and lat.(−46 °)–(−52 °) long. 360 ° − 60 °. Additional maximums diffusion occurs inside radiation belt. Appearance of these maximums seems to be closely connected with preceding powerful solar proton events and associated geomagnetic dynamics of new belt disturbances. After the series of solar proton events in June 1991 we observed significant enhancement of this new radiation belt formation. To achieve sufficient accuracy of dose rate predictions in low Earth orbits the structure and dynamics of new belt should be carefully analyzed to be included in a new environment model.

From the inter comparison of the data from “Liulin” and French developed tissue equivalent LET spectrometer NAUSICAA in the time period August–November 1992 we come to the following conclusions: Mainly there is good agreement between both data sets for absorbed dose in the region of SAA; Different situation of the instruments on the station can explain the cases when differences up to 2 times are observed; At high latitudes usually the tissue equivalent absorbed dose observations are 2 times larger than “Liulin” doses.     

Planetary Defense From Space: Part 1-Keplerian Theory

A system of two space bases housing missiles is proposed to achieve the Planetary Defense of the Earth against dangerous asteroids and comets. We show that the layout of the Earth–Moon system with the five relevant Lagrangian (or libration) points in space leads naturally to only one, unmistakable location of these two space bases within the sphere of influence of the Earth. These locations are at the two Lagrangian points L1 (in between the Earth and the Moon) and L3 (in the direction opposite to the Moon from the Earth).

We show that placing bases of missiles at L1 and L3 would cause those missiles to deflect the trajectory of asteroids by hitting them orthogonally to their impact trajectory toward the Earth, so as to maximize their deflection. We show that the confocal conics are the best class of trajectories fulfilling this orthogonal deflection requirement.

An additional remark is that the theory developed in this paper is just a beginning of a larger set of future research work. In fact, while in this paper we only develop the Keplerian analytical theory of the Optimal Planetary Defense achievable from the Earth–Moon Lagrangian points L1 and L3, much more sophisticated analytical refinements would be needed to:

1. Take into account many perturbation forces of all kinds acting on both the asteroids and missiles shot from L1 and L3;

2. add more (non-optimal) trajectories of missiles shot from either the Lagrangian points L4 and L5 of the Earth–Moon system or from the surface of the Moon itself;

3. encompass the full range of missiles currently available to the US (and possibly other countries) so as to really see “which asteroids could be diverted by which missiles”, even in the very simplified scheme outlined here.

Outlined for the first time in February 2002, our Confocal Planetary Defense concept is a Keplerian Theory that proved simple enough to catch the attention of scholars, representatives of the US Military and popular writers. These developments could possibly mark the beginning of an “all embracing” mathematical vision of Planetary Defense beyond all learned activities, dramatic movies and unknown military plans covered by secret.     

G-300, the first French Getaway Special microgravity measurements of fluid thermal conductivity

On Earth thermal conductivity measurements on liquids are difficult to perform because thermal motions due to convection. In microgravity the convection due to buoyancy is evanescent and we expect a strong lowering of Rayleigh and Nusselt numbers. Three low viscosity liquids are selected to carry out the measurements: distilled water (standard) and two silicone oils. We use a modified “hot plate” method with a simplified guard ring, the lowering of convective motions let us to use in the experimental cells larger interplate distances and/or temperature differences than in earth measurements so the accuracy must be improved. Comparisons between Earth and orbit results may help to understand the convection occurrence in our cells. G-300 payload is cantilevered from the experiment mounting plate (EMP) and it includes: four struts, an intermediate plate, a bottom plate with four bumpers, a battery box, an electronic rack and six experimental cells assembled in twin-packs thermally coupled with the EMP. Thermal, vibration and EMI tests have proved that the design satisfies the Nasa requirements and acceptance is in progress.     

Lunar base development missions

On 20 July 1969, humankind first set foot on our Moon. Since then we have developed the Space Shuttle, explored most of the planets, cooperated in the development of the International Space Station, and expanded our knowledge of the universe through use of systems such as the Hubble Space Telescope and the Mars Pathfinder. After just five human follow-on missions to our Moon, we have returned robotically only twice to orbit, to map the surface and explore for resources.

The indication of the presence of hydrogen concentration at the poles of our Moon found by Lunar Prospector has added a new perspective for groups studying and implementing future lunar missions. Plans for nearterm missions such as the European Space Agency (ESA) “Euromoon 2000”, the Japanese Lunar A and Selene, and the Mitsubishi ”Earthrise 2001” Project, along with follow-on phases to the Lunar Prospector, are the beginning of humankind's return to the Moon. Organizations such as the International Academy of Astronautics have long championed the “Case for an International Lunar Base,” and a vision of a commercially-based lunar program has been outlined by several groups. A Lunar Economic Development Authority (LEDA) promoted by the United Society in Space was promulgated by the filing of articles of incorporation in the state of Colorado on 4 August 1997. This non-profit corporation has as its goal the orderly development of the Moon, through issuance of bonds to international private citizens and business entities who care to invest in its long-term development.

This paper draws from the works of the aforementioned, and specifically from the International Academy of Astronautics Lunar Base Committee, to structure a series of architectures leading toward eventual international commercial colonization of the lunar surface. While the prospect of fully reusable transportation systems utilizing fully developed lunar resources to perpetuate the permanent lunar infrastructure is enticing, this is a goal. We must utilize our current and near-term capabilities to re-initiate human lunar presence, and then build on emerging technologies to strengthen our capabilities. Humankind's return to the Moon is a part of our destiny. We can return in the near future, and then proceed to a commercial, permanent settlement in the 21st century.     

Magnetobraking for Mars return vehicles

It is proposed that magnetobraking may be used to dissipate hyperbolic excess velocity from a spacecraft returning from Mars to Earth orbit. In magnetobraking, an electrodynamic tether is deployed from the spacecraft. The Earth's magnetic field produces a force on electrical current in the tether, which can be used to either brake or accelerate the spacecraft without expenditure of reaction mass. The peak acceleration on the Mars return is 0.007 m/s², and the amount of braking possible is dependent on the density and current-carrying capacity of the tether, but is independent of length. Since energy is produced as the spacecraft velocity decreases, no on-board power source is required. As the spacecraft approaches the Earth, the magnetic field increases and the power produced by the tether increases, reaching a maximum of about 800 W per kg of spacecraft mass at closest approach.     

Assessing the Risk of Orbital Debris Impact

A Space Debris Impact Risk Analysis Tool (SDIRAT) was developed and implemented to assess the orbital debris impact risk on a specified target in Earth orbit, in terms of flux, relative velocity, impact velocity, direction of the incoming particles, debris mass and diameter. Based on a deterministic approach, SDIRAT uses a realistic orbital debris population where each representative particle is identified by its rectangular coordinates (position and velocity) at a reference epoch. Using this information, some geometrical algorithms were developed and implemented to evaluate the contribution of each particle to the incoming flux. The position of the particle with respect to a specified target drives the selection criteria to reject, or select, it as a possible projectile. On the other hand, the relative velocity vector can be used to estimate the impact direction of the incoming flux. SDIRAT was conceived as a general tool for a variety of scenarios, such as low circular and elliptical orbits, up to the geosynchronous ring. This paper presents some examples of possible applications, including the computation of the incoming debris flux on SAX (low Earth orbit), SIRIO (geosynchronous orbit) and the IRIS upper stage (elliptical orbit). Other applications assess the impact risk for the Soviet Radar Ocean Reconnaissance Satellites Cosmos 1900 and Cosmos 1932.     

The earth explorer candidates: Nine focused missions, all satellite classes

The size of a satellite should follow naturally from the mission requirements taking into account implementation constraints, notably the choice of launcher and programmatics. The key concepts are “focused mission” and “maximum return for the investment”. The nine candidates proposed for the ESA Earth Explorer programme of research missions form a good representative set to illustrate a discussion of satellite classes. These focused missions lead to satellites from less than 100 kg and 100 W, to more than 2000 kg and 2 kW of power generation and include flights of opportunity and precursor experiments.     

Innovations for competitiveness: European views on “better-faster-cheaper”

The paper elaborates on “ lessons learned” from two recent ESA workshops, one focussing on the role of Innovation in the competitiveness of the space sector and the second on technology and engineering aspects conducive to better, faster and cheaper space programmes. The paper focuses primarily on four major aspects, namely:

1. a) the adaptations of industrial and public organisations to the global market needs;

2. b) the understanding of the bottleneck factors limiting competitiveness;

3. c) the trends toward new system architectures and new engineering and production methods;

4. d) the understanding of the role of new technology in the future applications.

Under the pressure of market forces and the influence of many global and regional players, applications of space systems and technology are becoming more and more competitive. It is well recognised that without major effort for innovation in industrial practices, organisations, R&D, marketing and financial approaches the European space sector will stagnate and loose its competence as well as its competitiveness. It is also recognised that a programme run according to the “better, faster, cheaper” philosophy relies on much closer integration of system design, development and verification, and draws heavily on a robust and comprehensive programme of technology development, which must run in parallel and off-line with respect to flight programmes.

A company's innovation capabilities will determine its future competitive advantage (in time, cost, performance or value) and overall growth potential. Innovation must be a process that can be counted on to provide repetitive, sustainable, long-term performance improvements. As such, it needs not depend on great breakthroughs in technology and concepts (which are accidental and rare). Rather, it could be based on bold evolution through the establishment of know-how, application of best practices, process effectiveness and high standards, performance measurement, and attention to customers and professional marketing. Having a technological lead allows industry to gain a competitive advantage in performance, cost and opportunities. Instrumental to better competitiveness is an R&D effort based on the adaptation of high technology products, capable of capturing new users, increasing production, decreasing the cost and delivery time and integrating high level of intelligence, information and autonomy. New systems will have to take in to account from the start what types of technologies are being developed or are already available in other areas outside space, and design their system accordingly. The future challenge for “faster, better, cheaper” appears to concern primarily “cost-effective”, performant autonomous spacecraft, “cost-effective”, reliable launching means and intelligent data fusion technologies and robust software serving mass- market real time services, distributed via EHF bands and Internet.

In conclusion, it can be noticed that in the past few years new approaches have considerably enlarged the ways in which space missions can be implemented. They are supported by true innovations in mission concepts, system architecture, development and technologies, in particular for the development of initiatives based on multi-mission mini-satellites platforms for communication and Earth observation missions. There are also definite limits to cost cutting (such as lowering heads counts and increasing efficiency), and therefore the strategic perspective must be shifted from the present emphasis on cost-driven enhancement to revenue-driven improvements for growth. And since the product life-cycle is continuously shortening, competitiveness is linked very strongly with the capability to generate new technology products which enhance cost/benefit performance.     

A permanently-acting NEA damage mitigation technique via the Yarkovsky effect

In the later stages of a combined mission of exploration and damage mitigation of a hazardous Near Earth Asteroid (NEA), sufficiently detailed tracking, orbit prediction, mass distribution, and physical properties data have been accumulated. It may be desirable to implement a “slow push” damage mitigation technique that is capable of altering the NEA orbit continuously and permanently. This possibility was suggested in a companion paper that outlined an archetypal exploration/mitigation mission to the NEA 99942 Apophis, which is achieved using a novel albedo change approach. The details and apparatus used in the albedo modification technique are described in the paper. We describe the design details and the constraints on particle size (to prevent electrostatic levitation and escape) and on the dispensing speed (to achieve the desired coverage zone and prevent particles from orbiting or escaping).     

Optimal control of a spinning double-pyramid Earth-pointing tethered formation

The dynamics and control of a tethered satellite formation for Earth-pointing observation missions is considered. For most practical applications in Earth orbit, a tether formation must be spinning in order to maintain tension in the tethers. It is possible to obtain periodic spinning solutions for a triangular formation whose initial conditions are close to the orbit normal. However, these solutions contain significant deviations of the satellites on a sphere relative to the desired Earth-pointing configuration. To maintain a plane of satellites spinning normal to the orbit plane, it is necessary to utilize “anchors”. Such a configuration resembles a double-pyramid. In this paper, control of a double-pyramid tethered formation is studied. The equations of motion are derived in a floating orbital coordinate system for the general case of an elliptic reference orbit. The motion of the satellites is derived assuming inelastic tethers that can vary in length in a controlled manner. Cartesian coordinates in a rotating reference frame attached to the desired spin frame provide a simple means of expressing the equations of motion, together with a set of constraint equations for the tether tensions. Periodic optimal control theory is applied to the system to determine sets of controlled periodic trajectories by varying the lengths of all interconnecting tethers (nine in total), as well as retrieval and simple reconfiguration trajectories. A modal analysis of the system is also performed using a lumped mass representation of the tethers.