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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 NTRS2017

Application of GPS to Enable Launch Vehicle Upper Stage Heliocentric Disposal

To properly dispose of the upper stage of the Space Launch System, the vehicle must perform a burn in Earth orbit to perform a close flyby of the Lunar surface to gain adequate energy to enter into heliocentric space. This architecture was selected to meet NASA requirements to limit orbital debris in the Earth-Moon system. The choice of a flyby for heliocentric disposal was driven by mission and vehicle constraints. This paper describes the SLS mission for Exploration Mission -1, a high level overview of the Block 1 vehicle, and the various disposal options considered. The research focuses on this analysis in terms of the mission design and navigation problem, focusing on the vehicle-level requirements that enable a successful mission. An inertial-only system is shown to be insufficient for heliocentric flyby due to large inertial integration errors from launch through disposal maneuver while on a trans-lunar trajectory. The various options for aiding the navigation system are presented and details are provided on the use of GPS to bound the state errors in orbit to improve the capability for stage disposal. The state estimation algorithm used is described as well as its capability in determination of the vehicle state at the start of the planned maneuver. This data, both dispersions on state and on errors, is then used to develop orbital targets to use for meeting the required Lunar flyby for entering onto a heliocentric trajectory. The effect of guidance and navigation errors on this capability is described as well as the identified constraints for achieving the disposal requirements. Additionally, discussion is provided on continued analysis and identification of system considerations that can drive the ability to integrate onto a vehicle intended for deep space.

Anzalone, Evan J.↗
NASA NTRS2020

NASA’s Space Launch System: Launch Capability for Lunar Exploration and Transformative Science

Excitement is building for the first launch of NASA’s Space Launch System (SLS), a unique exploration asset for the agency’s Artemis lunar program as well as for a new generation of science missions. SLS is designed for an array of missions beyond Earth’s orbit. The flexible system, which can be configured for Orion, cargo or Orion with co-manifested payload missions, offers high escape velocities to send more mass to deep space destinations. When configured with an 8.4 m-diameter fairing, SLS offers unmatched payload volume for human exploration and science missions. The initial Block 1 variant will insert at least 26 metric tons (t) to trans-lunar injection (TLI) and the more powerful Block 1B vehicle will launch 34-37 t to TLI using a new-development upper stage. Much of the initial SLS Block 1 vehicle is complete, including the upper stage and payload section, the core stage, engines and the solid rocket boosters. The first mission, Artemis I, launching from modernized and upgraded facilities at Kennedy Space Center (KSC), will be an uncrewed test flight of SLS, Orion and ground processing, with a primary objective of testing Orion’s heat shield at lunar re-entry velocity. Artemis I will have accommodations for 13 6U CubeSat payloads. These CubeSat missions will be deployed along the upper stage disposal trajectory after Orion separates from the vehicle. A rare opportunity for CubeSats to be deployed beyond low Earth orbit (LEO), Artemis I CubeSat missions range from searching for hydrogen and other volatiles on the lunar South Pole to studying the acceleration mechanisms of solar and interplanetary particles from a heliocentric trajectory. With manufacturing of the initial vehicle complete, fabrication and procurement is progressing for the second flight of SLS and Orion, Artemis II. Also an SLS Block 1 and Orion flight launching from KSC, Artemis II will mark the return of American astronauts to deep space with a lunar flyby-free return trajectory mission. With the Artemis III flight, NASA has the goal to land the first woman and the next man on the Moon. Infrastructure beyond SLS will be required for this effort, including elements of the lunar Gateway as well as lunar rovers, landers and additional commercially supplied launch services. SLS, as the only vehicle with the capability to lift 26 t of mass to TLI in its initial Block 1 variant, will remain a key component of this new-era exploration program. Future variants – Block 1B and Block 2 –will lift 34-45 t to TLI. This paper will discuss the status of testing and integration for the Artemis I vehicle, manufacturing progress for the second vehicle and the manifest outlook for primary, co-manifested and secondary payloads in the current deep space exploration environment.

Creech, Stephen D.↗
NASA NTRS1977

The Luna and Zond programs

These series have included flyby, lunar-orbiting, and soft-landing missions. A variety of experiments were carried out by these spacecraft including studies of magnetism, X-ray and gamma ray emissions, gravitational anomalies, and chemical composition. Soil samples, near- and farside photography (both color and black and white), and earth-cloud photography were also acquired.

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NASA NTRS1984

The Planetary Observer Program

An overview is presented of NASA's plans for the Planetary Observer Program, whose key element is to control the cost of each mission while establishing a long-term, stable base for the planetary sciences. The SSEC (Solar System Exploration Committee) has endorsed the view that many high science priority inner solar system missions are possible through the use of spacecraft derived from existing earth-orbital spacecraft. It has also recommended the application of space hardware such as that used on the Voyager and Galileo missions, development of both a new modular spacecraft for outer planet, comet, and main-belt asteroid missions (Mariner Mark II Program), and a multi-mission operations system to support future missions after the Venus Radar Mapper (VRM), and Galileo. A set of missions for the SSEC's Core Program has been recommended; they include: the VRM, the Mars Geoscience/Climatology Observer, the Comet Rendezvous/Asteroid Flyby, the Lunar Geoscience Orbiter, the Near-Earth Asteroid Rendezvous, the Venus Atmosphere Probe, the Mars Aeronomy Orbiter, the Mars Surface Probe, and the Comet Intercept Sample Return.

Blume, W. H.↗
NASA NTRS2014

Nuclear Thermal Propulsion: Past, Present, and a Look Ahead

NTR: High thrust high specific impulse (2 x LOXLH2 chemical) engine uses high power density fission reactor with enriched uranium fuel as thermal power source. Reactor heat is removed using H2 propellant which is then exhausted to produce thrust. Conventional chemical engine LH2 tanks, turbo pumps, regenerative nozzles and radiation-cooled shirt extensions used -- NTR is next evolutionary step in high performance liquid rocket engines.

Nuclear rocket engines↗
NASA NTRS2015

Affordable Development and Demonstration of a Small NTR engine and Stage: A Preliminary NASA, DOE, and Industry Assessment

In FY'11, Nuclear Thermal Propulsion (NTP) was identified as a key propulsion option under the Advanced In-Space Propulsion (AISP) component of NASA's Exploration Technology Development and Demonstration (ETDD) program A strategy was outlined by GRC and NASA HQ that included 2 key elements -"Foundational Technology Development" followed by specific "Technology Demonstration" projects. The "Technology Demonstration "element proposed ground technology demonstration (GTD) testing in the early 2020's, followed by a flight technology demonstration (FTD) mission by approx. 2025. In order to reduce development costs, the demonstration projects would focus on developing a small, low thrust (approx. 7.5 -16.5 klb(f)) engine that utilizes a "common" fuel element design scalable to the higher thrust (approx. 25 klb(f)) engines used in NASA's Mars DRA 5.0 study(NASA-SP-2009-566). Besides reducing development costs and allowing utilization of existing, flight proven engine hard-ware (e.g., hydrogen pumps and nozzles), small, lower thrust ground and flight demonstration engines can validate the technology and offer improved capability -increased payloads and decreased transit times -valued for robotic science missions identified in NASA's Decadal Study.

Spacecraft design↗
NASA NTRS2015

Affordable Development and Demonstration of a Small NTR Engine and Stage: How Small is Big Enough?

In FY11, NASA formulated a plan for Nuclear Thermal Propulsion (NTP) development that included Foundational Technology Development followed by system-level Technology Demonstrations The ongoing NTP project, funded by NASAs Advanced Exploration Systems (AES) program, is focused on Foundational Technology Development and includes 5 key task activities:(1) Fuel element fabrication and non-nuclear validation testing of heritage fuel options;(2) Engine conceptual design;(3) Mission analysis and engine requirements definition;(4) Identification of affordable options for ground testing; and(5) Formulation of an affordable and sustainable NTP development program Performance parameters for Point of Departure designs for a small criticality-limited and full size 25 klbf-class engine were developed during FYs 13-14 using heritage fuel element designs for both RoverNERVA Graphite Composite (GC) and Ceramic Metal (Cermet) fuel forms To focus the fuel development effort and maximize use of its resources, the AES program decided, in FY14, that a leader-follower down selection between GC and cermet fuel was required An Independent Review Panel (IRP) was convened by NASA and tasked with reviewing the available fuel data and making a recommendation to NASA. In February 2015, the IRP recommended and the AES program endorsed GC as the leader fuel In FY14, a preliminary development schedule DDTE plan was produced by GRC, DOE industry for the AES program. Assumptions, considerations and key task activities are presented here Two small (7.5 and 16.5 klbf) engine sizes were considered for ground and flight technology demonstration within a 10-year timeframe; their ability to support future human exploration missions was also examined and a recommendation on a preferred size is provided.

spacecraft design↗
NASA NTRS2016

Synchronized Lunar Pole Impact Plume Sample Return Trajectory Design

The presented trajectory design enables two maneuverable spacecraft launched onto the same trans-lunar injection trajectory to coordinate a steep impact of a lunar pole and subsequent sample return of the ejecta plume to Earth. To demonstrate this concept, the impactor is assumed to use the LCROSS missions trajectory and spacecraft architecture, thus the permanently-shadowed Cabeus crater on the lunar south pole is assumed as the impact site. The sample-return spacecraft is assumed to be a CubeSat that requires a complimentary trajectory design that avoids lunar impact after passing through the ejecta plume to enable sample-return to Earth via atmospheric entry.

impact plume↗
NASA NTRS2015

Affordable Development and Demonstration of a Small NTR Engine and Stage: A Preliminary NASA, DOE, and Industry Assessment

In FY11, NASA formulated a plan for Nuclear Thermal Propulsion (NTP) development that included Foundational Technology Development followed by system-level Technology Demonstrations. The ongoing NTP project, funded by NASA's Advanced Exploration Systems (AES) program, is focused on Foundational Technology Development and includes 5 key task activities: (1) Fuel element fabrication and non-nuclear validation testing of heritage fuel options; (2) Engine conceptual design; (3) Mission analysis and engine requirements definition; (4) Identification of affordable options for ground testing; and (5) Formulation of an affordable and sustainable NTP development program. Performance parameters for Point of Departure designs for a small criticality-limited and full size 25 klb (sub f) (kip, or 1000 pound-force)-class engine were developed during FYs 13-14 using heritage fuel element designs for both Rover NERVA (Nuclear Engine for Rocket Vehicle Applications) Graphite Composite (GC) and Ceramic Metal (Cermet) fuel forms. To focus the fuel development effort and maximize use of its resources, the AES program decided, in FY14, that a leader-follower down-selection between GC and cermet fuel was required. An Independent Review Panel (IRP) was convened by NASA and tasked with reviewing the available fuel data and making a recommendation to NASA. In February 2015, the IRP recommended and the AES program endorsed GC as the leader fuel In FY14, a preliminary development schedule DDTE (Design Development Test and Evaluation) plan was produced by GRC (Glenn Research Center), DOE (Department of Energy) and industry for the AES program. Assumptions, considerations and key task activities are presented here. At the direction of NASA HQ (32515), NASA and DOE are to work together to formulate a detailed development plan and schedule allowing the affordable development of a small (7.5-16.5 klbf) GC engine for possible flight technology demonstration (FTD) mission within a 10-year timeframe.

Lunar Flyby Mission↗
NASA NTRS2015

Exploring Our Solar System with CubeSats and NanoSats

The Jet Propulsion Laboratory (JPL) is NASA's lead center for robotic exploration of our solar system. We are known for our large, flagship missions, such as Voyager, which gave humanity its first close look at Jupiter and Saturn; and the Mars Rovers, which have excited millions worldwide with their daring landing exploits. Less familiar to those outside NASA may be our role in developing the Kepler mission, which has discovered more than 2000 planets around other stars; or the recently launched Soil Moisture Active Passive (SMAP) mission, one of many JPL Earth Science missions. A recent JPL initiative has emphasized low cost missions that use rapidly evolving technology developed for CubeSats and NanoSat s to explore our solar system. Costs are significantly lower (by one or two orders of magnitude) than for conventional JPL missions, and development time is also significantly shorter. At present 21 such CubeSat flight projects are under way at the laboratory with various partners : some in flight, some in development, some in advanced formulation. Four are planned as deep space missions. To succeed in exploring deep space CubeSat/NanoSat missions have to address several challenges: the more severe radiation environment, communications and navigation at a distance, propulsion, and packaging of instruments that can return valuable science into a compact volume/mass envelope. Instrument technologies, including cameras, magnetometers, spectrometers, radiometers, and even radars are undergoing miniaturization to fit on these smaller platforms. Other key technologies are being matured for smallsats and NanoSats in deep space, including micro -electric propulsion, compact radio (and optical) communications, and onboard data reduction. This paper will describe missions that utilize these developments including the first two deep space CubeSats (INSPIRE), planned for launch in 2017; the first pair of CubeSats to be sent to another planet (MARCO), manifested with the InSight Mars lander launch in March of 2016; a helicopter "drone" on Mars to extend the reach of future rovers; plans for a Lunar Flashlight mission to shine a light on the permanently shadowed craters of the Moon's poles; a Near Earth Asteroid CubeSat missio n; and a CubeSat constellation to demonstrate time series measurements of storm systems on Earth. From these beginnings, the potential for CubeSats and NanoSats to add to our knowledge of the solar system could easily grow exponentially. Imagine if every deep space mission carried one or more CubeSats that could operate independently (even for a brief period) on arrival at their target body. At only incremental additional cost, such spacecraft could go closer, probe deeper, and provide science measurements that we would not risk with the host spacecraft. This paper will describe examples including a NanoSat to probe the composition of Venus' atmosphere, impactors and close flybys of Europa, lunar probes, and soft landers for the moons of Mars. Low cost access to deep space also offers the potential for independent CubeSat/NanoSat missions - allowing us to characterize the population of near Earth asteroids for example, deploy a constellation around Venus, or take closer looks at the asteroid belt.

Deep Space↗
NASA NTRS2021

Effect of V-infinity Leveraging with Lunar-earth Gravity Assist on Interplanetary Trajectories

This work evaluates the effects of Lunar-Earth Gravity Assist on mission design and DV budget for interplanetary trajectories using Earth flybys and V1 Leveraging Maneuvers. The use of the Moon flyby, in combination to the Earth flyby, provides additional deviation of the hyperbolic excess velocity with respect to the Earth, also modifying its magnitude, and hence increases the overall flyby performance. The Lunar-Earth Gravity Assist has been evaluated for Exterior and Interior Leveraging, as well as for Short and Long transfers. A sample Jupiter mission is showcased, and the results suggest that the Moon effect during the flyby can reduce the total DVtot by 6.4% and the leveraging DVVILM by 21.06%.

Campagnola, Stefano↗
NASA NTRS1992

Galileo Earth Moon Flyby

This video has five sections. The first is a live discussion of the information that scientists hope to gain by the Galileo flyby of the Moon. This section has no introduction. There is a great deal of the discussion about the lunar craters and lunar volcanism. There is also some discussion of the composition of the far side of the moon. The second section is a short animation that shows the final step to Jupiter with particular emphasis on the gravitational assisted velocity boost, which was planned to give the spacecraft the requisite velocity to make the trip to Jupiter. The next section is an update of the status of the flyby of the Moon, and the Earth, with an explanation of the trajectory around the earth, and the moon. A photograph of the tracking station in Canberra, Australia is included. The next section is a tour of a full-scale model of the spacecraft. The last section is a discussion with the person charged with the procurement of the instrumentation aboard the spacecraft; the importance of the lunar flyby to assist in the calibration of the instruments is discussed.

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NASA NTRS1993

Lunar multispectral mosaics from Galileo's second Earth-Moon flyby

Galileo's Solid-State Imaging (SSI) experiment acquired about 800 images of the Moon from the second Earth-Moon flyby (EM2) in December of 1992. Ten major sequences were acquired; each consists of mosaics of the entire or nearly entire visible and illuminated surface from each viewing geometry in at least six spectral filters (effective wavelengths for the Moon of 420, 564, 660, 756, 890, and 990 nm). The geometries of LUNMOS numbers 3, 4, 5, and 6 were designed to provide stereo data at the best possible resolutions. The purpose of this abstract is to describe the sequences, calibration, processing, and mosaicking, and to present a set of color products in a poster session.

Mcewen, A. S.↗
NASA NTRS2014

Asteroid Redirect Crewed Mission Nominal Design and Performance

In 2010, the President announced that, in 2025, the U.S. intended to launch a human mission to an asteroid [1]. This announcement was followed by the idea of a Capability Driven Framework (CDF) [2], which is based on the idea of evolving capabilities from less demanding to more demanding missions to multiple possible destinations and with increased flexibility, cost effectiveness and sustainability. Focused missions, such as a NASA inter-Center study that examined the viability and implications of sending a crew to a Near Earth Asteroid (NEA) [3], provided a way to better understand and evaluate the utility of these CDF capabilities when applied to an actual mission. The long duration of the NEA missions were contrasted with a concept described in a study prepared for the Keck Institute of Space Studies (KISS) [4] where a robotic spacecraft would redirect an asteroid to the Earth-Moon vicinity, where a relatively short duration crewed mission could be conducted to the captured asteroid. This mission concept was included in the National Aeronautics and Space Administration (NASA) fiscal year 2014 budget request, as submitted by the NASA Administrator [5]. NASA studies continued to examine the idea of a crewed mission to a captured asteroid in the Earth-Moon vicinity. During this time was an announcement of NASA's Asteroid Grand Challenge [6]. Key goals for the Asteroid Grand Challenge are to locate, redirect, and explore an asteroid, as well as find and plan for asteroid threats. An Asteroid Redirect Mission (ARM) study was being conducted, which supports this Grand Challenge by providing understanding in how to execute an asteroid rendezvous, capture it, and redirect it to Earth-Moon space, and, in particular, to a distant retrograde orbit (DRO). Subsequent to the returning of the asteroid to a DRO, would be the launch of a crewed mission to rendezvous with the redirected asteroid. This report examines that crewed mission by assessing the Asteroid Redirect Crewed Mission (ARCM) nominal design and performance costs associated with an Orion based crewed rendezvous mission to a captured asteroid in an Earth-Moon DRO. The ARM study includes two fundamental mission phases: 1) The Asteroid Redirect Robotic Mission (ARRM) and 2) the ARCM. The ARRM includes a solar electric propulsion based robotic asteroid return vehicle (ARV) sent to rendezvous with a selected near Earth asteroid, capture it, and return it to a DRO in the Earth-Moon vicinity. The DRO is selected over other possible asteroid parking orbits due to its achievability (by both the robotic and crewed vehicles) and by its stability (e.g., no orbit maintenance is required). After the return of the asteroid to the Earth-Moon vicinity, the ARCM is executed and carries a crew of two astronauts to a DRO to rendezvous with the awaiting ARV with the asteroid. The outbound and inbound transfers employ lunar gravity assist (LGA) flybys to reduce the Orion propellant requirement for the overall nominal mission, which provides a nominal mission with some reserve propellant for possible abort situations. The nominal mission described in this report provides a better understanding of the mission considerations as well as the feasibility of such a crewed mission, particularly with regard to spacecraft currently undergoing development, such as the Orion vehicle and the Space Launch System (SLS).

Condon, Gerald↗
NASA NTRS1992

Use of multiple lunar swingby for departure to Mars

Single and double lunar swingbys have previously been suggested for earth escape missions. This paper investigates the payload gain available from single and double lunar gravity assists. The gain available from a powered earth flyby used in conjunction with a double lunar gravity assist is also discussed. This paper also examines return to the earth-moon system for a third lunar gravity assist and the payload gain that results. Use of these methods allows a spacecraft to depart earth with an energy less than that needed to escape and to proceed on to planetary encounters. Opportunities for travel to Mars using these methods are presented. Fairly regular opportunities exist that result in payload gain. The main problem is that launch windows are rather restrictive. Mission scenarios where this procedure may be useful are presented.

Hanson, John M.↗
NASA NTRS1991

Space transfer vehicle concepts and requirements. Volume 1: Executive summary

The objectives of the Space Transfer Vehicle (STV) Concepts and Requirements studies were to provide sensitivity data on usage, economics, and technology associated with new space transportation systems. The study was structured to utilize data on the emerging launch vehicles, the latest mission scenarios, and Space Exploration Initiative (SEI) payload manifesting and schedules, to define a flexible, high performance, cost effective, evolutionary space transportation system for NASA. Initial activities were to support the MSFC effort in the preparation of inputs to the 90 Day Report to the National Space Council (NSC). With the results of this study establishing a point-of-departure for continuing the STV studies in 1990, additional options and mission architectures were defined. The continuing studies will update and expand the parametrics, assess new cargo and manned ETO vehicles, determine impacts on the redefined Phase 0 Space Station Freedom, and to develop a design that encompasses adequate configuration flexibility to ensure compliance with on-going NASA study recommendations with major system disconnects. In terms of general requirements, the objectives of the STV system and its mission profiles will address crew safety and mission success through a failure-tolerant and forgiving design approach. These objectives were addressed through the following: engine-out capability for all mission phases; built-in-test for vehicle health monitoring to allow testing of all critical functions such as, verification of lunar landing and ascent engines before initiating the landing sequence; critical subsystems will have multiple strings for redundancy plus adequate supplies of onboard spares for removal and replacement of failed items; crew radiation protection; and trajectories that optimize lunar and Mars performance and flyby abort capabilities.

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