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Search indexed NASA NTRS and DOE OSTI research on propulsion, heat transfer, battery materials and energy systems. Follow report and document links to the original sources.

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NASA NTRS2019

NASA's Space Launch System: Payload Opportunities for Lunar Exploration, Science Missions

Nearing its first launch, the Space Launch System (SLS), NASA’s new super heavy-lift launch vehicle, offers highly energetic launches that deliver more mass and provide more volume in 8.4 m-diameter and potentially larger fairings to make a new generation of deep space missions possible. An evolvable launcher available in crew, crew with a co-manifested payload (CPL) and cargo-only configurations, SLS is a crucial capability to enable astronauts to return to the Moon, but it also offers key benefits for science missions. NASA’s 21st-century return to the Moon recently received a formal name: the Artemis program. In addition to the core enabling capabilities of SLS and the Orion crew spacecraft, NASA will also enlist international and commercial partnerships for Artemis. The Agency intends to build a scientific outpost in lunar orbit, the Gateway, from which human and robotic missions to and from the surface can rendezvous. SLS will launch Orion on a series of missions leading to landing the first woman and the next man on the Moon as part of Artemis. SLS uses proven propulsion systems: two solid rocket boosters and four RS-25s engines that have been upgraded to provide more thrust and operate in the SLS environment. SLS Block 1 uses a modified Delta IV Heavy upper stage, called the Interim Cryogenic Propulsion Stage (ICPS) and lifts at least 26 metric tons (t) to trans-lunar injection (TLI). The Block 1 vehicle can also be outfitted with a 5 m-diameter fairing. Block 1B, the next major variant, also uses solid rocket boosters and RS-25 engines to achieve Earth escape velocity, but replaces the single-engine liquid hydrogen (LH2)/liquid oxygen (LOX) ICPS with a four-engine LH2/LOX Exploration Upper Stage (EUS) to increase mass to TLI to 34-37 t, depending on crew or cargo configuration. In the Block 1B crew configuration, a 10 m-tall Universal Stage Adapter (USA) connects the vehicle to Orion and can carry a CPL up to 10 t. The USA provides 286 m3 of unpressurized volume for payloads. For large payloads, 8.4 m- and 10 m-diameter cargo fairings in 19.1 m and 27.4 m lengths are possible. The ultimate SLS vehicle, Block 2, incorporates evolved boosters to reach a lift capacity of more than 45 t to TLI. The capabilities of SLS not only make new missions to the Moon possible, but also game-changing science missions, such as deployment of large-aperture space telescopes, spacecraft to the ice giants or even probes to the interstellar medium. This paper will discuss the capabilities of SLS, the vehicle’s planned evolution, missions that can effectively utilize the vehicle and manufacturing status of the vehicle.

Creech, Steve↗
NASA NTRS2009

T-Rex: A Japanese Space Tether Experiment

Electrodynamic tether (EDT) thrusters work by virtue of the force a magnetic field exerts on a wire carrying an electrical current. The force, which acts on any charged particle moving through a magnetic field (including the electrons moving in a current-carrying wire), were concisely expressed by Lorentz in 1895 in an equation that now bears his name. The force acts in a direction perpendicular to both the direction of current flow and the magnetic field vector. Electric motors make use of this force: a wire loop in a magnetic field is made to rotate by the torque the Lorentz Force exerts on it due to an alternating current in the loop times so as to keep the torque acting in the same sense. The motion of the loop is transmitted to a shaft, thus providing work. Although the working principle of EDT thrusters is not new, its application to space transportation may be significant. In essence, an EDT thruster is just a clever way of getting an electrical current to flow in a long orbiting wire (the tether) so that the Earth s magnetic field will accelerate the wire and, consequently the payload attached to the wire. The direction of current flow in the tether, either toward or away from the Earth along the local vertical, determines whether the magnetic force will raise or lower the orbit. The bias voltage of a vertically deployed metal tether, which results just from its orbital motion (assumed eastward) through Earth s magnetic field, is positive with respect to the ambient plasma at the top and negative at the bottom. This polarization is due to the action of the Lorentz force on the electrons in the tether. Thus, the natural current flow is the result of negative electrons being attracted to the upper end and then returned to the plasma at the lower end. The magnetic force in this case has a component opposite to the direction of motion, and thus leads to a lowering of the orbit and eventually to re-entry. In this generator mode of operation the Lorentz Force serves both to drive the current and then to act on the current to decelerate the system. One of the most important features of tether thrusters is that they use renewable energy sources to drive the electrical current flow in either the orbit-raising or orbit-lowering modes. Sources inherent to Earth orbit are used. To raise the orbit, sunlight can be converted to the electrical energy required to drive the tether current. To lower the orbit, the orbital energy itself (supplied by the Earth-to-orbit launcher when it raises the system into orbit) is the energy source of the tether current via the action of the Lorentz Force. Electrodynamic tethers can be directly applied to a wide spectrum of uses in space. As a propulsion system, they include satellite de-orbit, transfer of a satellite from one orbit to another, altitude maintenance for large spacecraft such as the International Space Station, and since it works wherever there is a magnetic field and an ionosphere planetary exploration missions. An electrodynamic tether upper stage could be used as an Orbit Transfer Vehicle (OTV) to move payloads within low earth orbit. The OTV would rendezvous with the payload and launch vehicle, grapple the payload and maneuver it to a new orbital altitude or inclination without the use of boost propellant. The tug could then lower its orbit to rendezvous with the next payload and repeat the process. Conceivably, such a system could perform several orbital maneuvering assignments without resupply, making it relatively inexpensive to operate.

Johnson, Les↗
NASA NTRS1982

Design of a 7kW power transfer solar array drive mechanism

With the availability of the Shuttle and the European launcher, Ariane, there will be a continuing trend towards large payload satellite missions requiring high-power, high-inertia, flexible solar arrays. The need arises for a solar array drive with a large power transfer capability which can rotate these solar arrays without disturbing the satellite body pointing. The modular design of such a Solar Array Drive Mechanism (SADM) which is capable of transferring 7kW of power or more is described. Total design flexibility has been achieved, enabling different spacecraft power requirements to be accommodated within the SADM design.

Sheppard, J. G.↗
NASA NTRS2020

PHALANX: Expendable Projectile Sensor Networks for Planetary Exploration

Technologies enabling long-term, wide-ranging measurement in hard-to-reach areas are a critical need for planetary science inquiry. Phenomena of interest include flows or variations in volatiles, gas composition or concentration, particulate density, or even simply temperature. Improved measurement of these processes enables understanding of exotic geologies and distributions or correlating indicators of trapped water or biological activity. However, such data is often needed in unsafe areas such as caves, lava tubes, or steep ravines not easily reached by current spacecraft and planetary robots. To address this capability gap, we have developed miniaturized, expendable sensors which can be ballistically lobbed from a robotic rover or static lander - or even dropped during a flyover. These projectiles can perform sensing during flight and after anchoring to terrain features. By augmenting exploration systems with these sensors, we can extend situational awareness, perform long-duration monitoring, and reduce utilization of primary mobility resources, all of which are crucial in surface missions. We call the integrated payload that includes a cold gas launcher, smart projectiles, planning software, network discovery, and science sensing: PHALANX. In this paper, we introduce the mission architecture for PHALANX and describe an exploration concept that pairs projectile sensors with a rover “mothership.” Science use cases explored include reconnaissance using ballistic cameras, volatiles detection, and building timelapse maps of temperature and illumination conditions. Strategies to autonomously coordinate constellations of deployed sensors to self-discover and localize with peer ranging (i.e. a “local GPS”) are summarized, thus providing communications infrastructure beyond-line-of-sight (BLOS) of the rover. Capabilities were demonstrated through both simulation and physical testing with a terrestrial prototype. The approach to developing a terrestrial prototype is discussed, including design of the launching mechanism, projectile optimization, micro-electronics fabrication, and sensor selection. Results from early testing and characterization of commercial-off-the-shelf (COTS) components are reported. Nodes were subjected to successful burn-in tests over 48 hours at full logging duty cycle. Integrated field tests were conducted in the Roverscape, a half-acre planetary analog environment at NASA Ames, where we tested up to 10 sensor nodes simultaneously coordinating with an exploration rover. Ranging accuracy has been demonstrated to be within +/-10cm over 20m using commodity radios when compared to high-resolution laser scanner ground truthing. Evolution of the design, including progressive miniaturization of the electronics and iterated modifications of the enclosure housing for streamlining and optimized radio performance are described. Finally, lessons learned to date, gaps toward eventual flight mission implementation, and continuing future development plans are discussed.

Dille, Michael↗
NASA NTRS1966

Postflight Evaluation of Atlas-Centaur AC-6 (Launched August 11, 1965)

The sixth Atlas Centaur vehicle (AC-6) was successfully launched from the Eastern T e s t Range, Complex 36B, on August 11, 1965, at 0931:04.430 EST. A 2084-pound dynamic model of the Surveyor payload was placed in a simulated lunar transfer trajectory. Vehicle systems operated satisfactorily and all the flight objectives were accomplished. Lift-off within 4 seconds of the window opening demonstrated the launch-on-time capability of the vehicle were accurately compensated for by the Centaur guidance system. the Surveyor model into a near-perfect lunar transfer trajectory would have resulted in an impact of the moon without a midcourse correction. To hit the precise target area on the lunar surface, the required correction would have been 4.25 meters per second, which is well within the spacecraft capability. Normal thrust and impulse levels were obtained with both the A t l a s and Centaur propulsion systems. However, a sizeable thrust overshoot on startup of the Centaur engines has not been resolved. A propellant-utilization system used for the first time on the Centaur, accurately controlled the fuel and oxidant consumption. The turnaround and retrothrust maneuver were performed without incident. Relatively high longitudinal modal excitations and lateral payload excitations were obtained at lift-off; these high perturbations are believed t o be related t o the launcher holddown arms. Nominal temperatures were recorded for both the external vehicle skin and the payload compartment; however, abnormally low temperatures were measured in the forward equipment area, which may have resulted from leakage of cold helium purge gas. All vehicle electrical systems performed satisfactorily; the only difficulty with the RF systems was obtained with the C-band transponder. of the vehicle instrumentation yielded valid data. The AC-6 vehicle was constructed with several new lightweight designs including the forward bulkhead, thrust barrel, interstage adapter and tank skin thickness reduction from 0.016 t o 0.014 inch. No deficiencies were observed in any of these new structural elements.

Source record↗
NASA NTRS1994

Minuteman 2 launched small satellite

The goal of LEOSat Industries' Spring 1994 project was to design a small satellite that has a strong technology demonstration or scientific justification and incorporates a high level of student involvement. The satellite is to be launched into low earth orbit by the converted Minuteman 2 satellite launcher designed by Minotaur Designs, Inc. in 1993. The launch vehicle shroud was modified to a height of 90 inches, a diameter of 48 inches at the bottom and 35 inches at the top for a total volume of 85 cubic feet. The maximum allowable mass of the payload is about 1100 lb., depending on the launch site, orbit altitude, and inclination. The satellite designed by LEOSat Industries is TerraSat, a remote-sensing satellite that will provide information for use in space-based earth studies. It will consist of infrared and ultraviolet/visible sensors similar to the SDI-developed sensors being tested on Clementine. The sensors will be mounted on the Defense Systems, Inc. Standard Satellite-1 spacecraft bus. LEOSat has planned for two satellites orbiting the Earth with trajectories similar to that of LANDSAT 5. The semi-major axis is 7080 kilometers, the eccentricity is 0, and the inclination is 98.2 degrees. The estimated mass of TerraSat is 145 kilograms and the estimated volume is 1.8 cubic meters. The estimated cost of TerraSat is $13.7 million. The projected length of time from assembly of the sensors to launch of the spacecraft is 13 months.

Chan, Sunny↗
NASA NTRS1992

Planet-B: A Japanese Mars aeronomy observer

An introduction is given to a Japanese Mars mission (Planet-B) which is being planned at the Institute of Space and Aeronautical Science (ISAS), Japan. Planet-B aims to study the upper atmosphere of Mars and its interaction with the solar wind. The launch of Planet-B is planned for 1996 on a new launcher, M-L, which is being developed at ISAS. In addition to the interaction with the solar wind, the structure of the Martian upper atmosphere is thought to be controlled by the meteorological condition in the lower atmosphere. The orbit of Planet-B was chosen so that it will pass two important regions, the region where the solar wind interacts with the Martian upper atmosphere and the tail region where ion acceleration is taking place. Considering the drag due to the Martian atmosphere, the periapsis altitude of 150 km and apoapsis of 10 Martian radii are planned. The orbit plane will be nearly parallel to the ecliptic plane. The altitude of the spacecraft will be spin stabilized and its spin axis will be controlled to the point of the earth. The dry weight of the spacecraft will be about 250 kg, including the scientific payload which consists of a magnetometer, plasma instruments, HF sounder, UV imaging spectrometer, and lower atmosphere monitor.

Tsuruda, K.↗
NASA NTRS2009

Ares V: Game Changer for National Security Launch

NASA is designing the Ares V cargo launch vehicle to vastly expand exploration of the Moon begun in the Apollo program and enable the exploration of Mars and beyond. As the largest launcher in history, Ares V also represents a national asset offering unprecedented opportunities for new science, national security, and commercial missions of unmatched size and scope. The Ares V is the heavy-lift component of NASA's dual-launch architecture that will replace the current space shuttle fleet, complete the International Space Station, and establish a permanent human presence on the Moon as a stepping-stone to destinations beyond. During extensive independent and internal architecture and vehicle trade studies as part of the Exploration Systems Architecture Study (ESAS), NASA selected the Ares I crew launch vehicle and the Ares V to support future exploration. The smaller Ares I will launch the Orion crew exploration vehicle with four to six astronauts into orbit. The Ares V is designed to carry the Altair lunar lander into orbit, rendezvous with Orion, and send the mated spacecraft toward lunar orbit. The Ares V will be the largest and most powerful launch vehicle in history, providing unprecedented payload mass and volume to establish a permanent lunar outpost and explore significantly more of the lunar surface than was done during the Apollo missions. The Ares V consists of a Core Stage, two Reusable Solid Rocket Boosters (RSRBs), Earth Departure Stage (EDS), and a payload shroud. For lunar missions, the shroud would cover the Lunar Surface Access Module (LSAM). The Ares V Core Stage is 33 feet in diameter and 212 feet in length, making it the largest rocket stage ever built. It is the same diameter as the Saturn V first stage, the S-IC. However, its length is about the same as the combined length of the Saturn V first and second stages. The Core Stage uses a cluster of five Pratt & Whitney Rocketdyne RS-68B rocket engines, each supplying about 700,000 pounds of thrust. Its propellants are liquid hydrogen and liquid oxygen. The two solid rocket boosters provide about 3.5 million pounds of thrust at liftoff. These 5.5-segment boosters are derived from the 4-segment boosters now used on the Space Shuttle, and are similar to those used in the Ares I first stage. The EDS is powered by one J-2X engine. The J-2X, which has roughly 294,000 pounds of thrust, also powers the Ares I Upper Stage. It is derived from the J-2 that powered the Saturn V second and third stages. The EDS performs two functions. Its initial suborbital burns will place the lunar lander into a stable Earth orbit. After the Orion crew vehicle, launched separately on an Ares I, docks with the lander/EDS stack, EDS will ignite a second time to put the combined 65-metric ton vehicle into a lunar transfer orbit. When it stands on the launch pad at Kennedy Space Center late in the next decade, the Ares V stack will be approximately 381 feet tall and have a gross liftoff mass of 8.1 million pounds. The current point-of-departure design exceeds Saturn V s mass capability by approximately 40 percent. Using the current payload shroud design, Ares V can carry 315,000 pounds to 29-degree low Earth orbit (LEO) or 77,000 pounds to a geosynchronous orbit. Another unique aspect of the Ares V is the 33-foot-diameter payload shroud, which encloses approximately 30,400 cubic feet of usable volume. A larger hypothetical shroud for encapsulating larger payloads has been studied. While Ares V makes possible larger payload masses and volumes, it may alternately make possible more cost-effective mission design if the relevant payload communities are willing to consider an alternative to the existing approach that has driven them to employ complexity to solve current launch vehicle mass and volume constraints. By using Ares V s mass and volume capabilities as margin, payload designers stand to reduce development risk and cost. Significant progress has been made on the Ares V to support a plaed fiscal 2011 authority-to-proceed (ATP) milestone. The Ares V team is actively reaching out to external organizations during this early concept phase to ensure that the Ares V vehicle can be leveraged for national security, science, and commercial development needs. This presentation will discuss Ares V vehicle configuration, the path to the current concept, accomplishments to date, and potential payload utilization opportunities.

Sumrall, Phil↗
NASA NTRSDate not supplied

SERENE: Saturn Enceladus Return Explorer with Nuclear Electric Propulsion

A ‘quick’ Enceladus sample return mission concept was developed based on the scientist recommendations at the recent ‘Accelerating Space Science with Nuclear Technology Workshop’. The Nuclear Electric Propulsion spacecraft assumed a follow-on 40 kWe nuclear reactor using the demonstrated 40 kWe Fission Surface Power system, expected in the early 2030s. The NEP vehicle also utilized a set of NEXT-C ion thrusters as well as planned Artemis commercial launchers. By launching the 40 kW NEP vehicle on a Starship and adding the propellants of 15 tankers, the 27t probe could be sent on a direct trajectory to Saturn (no Earth or Jupiter flybys) where NEP was used for Saturn capture, spiral down, spiral up and return to the Earth. A small lander obtained the surface Enceladus sample. Using the 40 kW NEP provided a round-trip time of only 16.5 years. A second option was more attractive from a science perspective whereby the NEP vehicle would deliver a large, 6t chemical lander to low Enceladus orbit where it would grab and return a sample to Earth (similar to the recent Orbilander design but in reverse). After deploying the lander, the NEP vehicle would stay in Saturn space performing a moon tour by orbiting four more moons and mapping the large moon of Titan. This option took slightly longer (18.5 yrs) due to the chemical return leg limitations. Both options demonstrated the agility, payload capability, and sample return goals the workshop recommended. An all-chemical option with two stages was roughly analyzed but took 21.5 yrs and required a Jupiter gravity assist.

Nuclear Electric Propulsion↗
NASA NTRS2018

Payload Utilization in NASA's Space Launch System

With Space Policy Directive 1, the United States administration has directed the National Aeronautics and Space Administration’s (NASA’s) Human Exploration & Operations Mission Directorate (HEOMD) to return to the Moon with missions and infrastructure designed to support a sustained presence in cislunar space, with robotic and human lunar surface operations. NASA’s new deep space exploration system — the super heavy-lift Space Launch System (SLS), the Orion crew spacecraft and revamped launch facilities at Kennedy Space Center (KSC) — will enable NASA and its commercial and international partners to meet this goal for human exploration of deep space. SLS is the most capable launch vehicle for these efforts, as well as for sending robotic missions deep into the solar system, or even to interstellar space. The vehicle will be available in crew and cargo configurations in progressively more powerful block variants. The initial Block 1 lift capability of at least 26 metric tons (t) to trans-lunar injection (TLI) will be followed by a more powerful Block 1B with the power to loft more than 37 t to TLI. The ultimate Block 2 variant will lift more than 45 t to TLI. For payload accommodation, the Block 1 vehicle can utilize a 5 meter (m) fairing in its cargo configuration with the crew version also able to provide berths for 6U and 12U CubeSats as secondary payloads. The Block 1B crew vehicle will provide as much volume as the space shuttle payload bay in a Universal Stage Adapter (USA) for co-manifested payloads (CPLs). Block 1B cargo vehicles will offer 8.4 m-diameter fairings in 19.1 m and possibly longer lengths, with enough volume to accommodate lunar-orbiting habitat modules and other elements of NASA’s Gateway science outpost. For Mars-class payloads, larger fairings for the Block 2 cargo launcher are under consideration. For missions beyond the Earth-Moon system, SLS offers greater characteristic energy (C3) than any other launch vehicle, enabling shorter transit times or heavier payloads with more robust science packages for missions to the outer solar system. Indeed, the unmatched combination of thrust, payload volume and departure energy that SLS provides opens new opportunities for human and robotic exploration of deep space. This paper will provide an overview of the various vehicle block configurations, their capabilities and payload accommodations for sending primary, co-manifested and secondary payloads to deep space

Creech, Stephen D.↗
NASA NTRS2019

NASA's Space Launch System: Deep Space Access for CubeSats

NASA is embarking on a new generation of missions to the Moon, known as the Artemis program. The Agency’s new super heavy-lift launch vehicle, the Space Launch System (SLS), is a critical enabling capability for these efforts, which will serve as a proving ground for future crewed expeditions to Mars. SLS is designed to return astronauts to the Moon in the Orion spacecraft and to launch more mass and provide more volume for critical payloads than commercially available vehicles, giving NASA a unique asset for deep space exploration. When performance margin and volume is available, as it is on the first flight, Artemis I (previously Exploration Mission-1), SLS can offer CubeSats access to deep space, beyond Earth’s orbit. The Artemis I flight has 13 6U (14 kg) CubeSats manifested, and the Program is currently accepting proposals for 6U and 12U payloads for the Artemis II flight through the agency’s CubeSat Launch Initiative (CSLI) program. Proposals from U.S.-based payload developers to fly on Artemis II will be accepted through CSLI until November 4, 2019. When ridesharing on SLS deep space missions, CubeSats must demonstrate they will not interfere with primary mission objectives. The 13 Artemis I CubeSats that hail from industry, academia, NASA and its international partners represent an array of exciting deep space science investigations and technology demonstrations that may help inform future Artemis missions. The initial SLS Block 1 vehicle for the Artemis I flight is fully manufactured; several elements are complete and have been delivered to the Exploration Ground Systems (EGS) Program at Kennedy Space Center (KSC), which has responsibility for integrating and launching the SLS and Orion stack. Completed elements of the Artemis I vehicle include the Orion Stage Adapter (OSA), which houses the 13 Artemis I CubeSats. With the Artemis I flight hardware and software nearing completion, work is in progress for the second Block 1 launcher, designated for the crewed Artemis II flight. Hardware is being manufactured for every element of the Artemis II vehicle. In this paper, the author will review the status of the Artemis I vehicle and the payloads manifested for the initial flight. Information on the Secondary Payload Deployment System (SPDS) that the SLS Program provides to payload developers will also be included.

Creech, Stephen D.↗
NASA NTRS2012

Advanced Lithium Ion Venus Explorer (ALIVE)

The COncurrent Multidisciplinary Preliminary Assessment of Space Systems (COMPASS) Team partnered with the Applied Research Laboratory to perform a NASA Innovative Advanced Concepts (NIAC) Program study to evaluate chemical based power systems for keeping a Venus lander alive (power and cooling) and functional for a period of days. The mission class targeted was either a Discovery ($500M) or New Frontiers ($750M to $780M) class mission. Historic Soviet Venus landers have only lasted on the order of 2 hours in the extreme Venus environment:temperatures of 460 degrees Centigrade and pressures of 93 bar. Longer duration missions have been studied using plutonium powered systems to operate and cool landers for up to a year. However, the plutonium load is very large. This NIAC study sought to still provide power and cooling but without the plutonium. Batteries are far too heavy but a system which uses the atmosphere (primarily carbon dioxide) and on on-board fuel to power a power generation and cooling system was sought. The resuling design was the Advanced Long-Life Lander Investigating the Venus Environment (ALIVE) Spacecraft (S/C) which burns lithium (Li) with the CO2 atmosphere to heat a Duplex Stirling to power and cool the lander for a 5-day duration (until the Li is exhausted). While it does not last years a chemical powered system surviving days eliminates the cost associated with utilizing a flyby relay S/C and allows a continuous low data rate direct to earth (DTE) link in this instance from the Ovda Regio of Venus. The five-day collection time provided by the chemical power systems also enables science personnel on earth to interact and retarget science - something not possible with an approximately 2-hour spacecraft lifetime. It also allows for contingency operations directed by the ground (reduced risk). The science package was based on that envisioned by the Venus Intrepid Tessera Lander (VITaL) Decadal Survey Study. The Li Burner within the long duration power system creates approximately 14000 W of heat. This 1300 degree Centigrade heat using Li in the bottom "ballast" tank is melted to liquid by the Venus temperature, drawn into a furnace by a wick and burned with atmospheric CO2. The Li carbonate exhaust is liquid at 1300 degrees Centigrade and being denser than Li drains into the the Li tank and solidifies. Since the exhaust product is a dense liquid no "chimney" is required which conserves the heat for the stirling power convertor. The Duplex Stirling provides about 300 W of power and removes about 300 W of heat from the avionics and heat that leaks into the 1-bar-insulated payload pressure vessel kept at 25 degrees Centigrade. The Na K radiator is run to the top of the drag flap.The ALIVE vehicle is carried to Venus via an Atlas 411 launch vehicle (LV) with a C3 of 7 km2/s2. An Aeroshell, derived from the Genesis mission, enables a direct entry into the atmosphere of Venus (-10 degrees Centigrade, 40 g max) and 6 m/s for landing (44 g) using a drag ring. For surface science and communication, a 100 WRF (WebEx Recording Format), X-Band 0.6-meter pointable DTE (Direct-to-Earth) antenna provides 2 kbps (kilobits per second) to DSN (Deep-Space Network) 34-meter antenna clusters.Table 1.1 summarizes the top-level details of each subsystem that was incorporated into the design. Cost estimates of the ALIVE mission show it at approximately $760M which puts it into the New Frontiers class.The ALIVE landed duration is only limited by the amount of Li which can be carried by the lander. Further studies are needed to investigate how additional mass can be carried, perhaps by a larger launcher and larger aeroshell.

Exploration↗