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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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All sourcesNASA NTRSDOE OSTI
NASA NTRS2017

The Lunar IceCube Mission Challenge: Attaining Science Orbit Parameters from a Constrained Approach Trajectory

The challenges of targeting specific lunar science orbit parameters from a concomitant Sun-EarthMoon system trajectory are examined. While the concept of ballistic lunar capture is well-studied, achieving and controlling the time evolution of the orbital elements to satisfy mission constraints is especially problematic when the spacecraft is equipped with a low-thrust propulsion system. Satisfying these requirements on the lunar approach and capture segments is critical to the success of the Lunar IceCube mission, a 6U CubeSat that will prospect for water in solid (ice), liquid, and vapor forms and other lunar volatiles from a low-periapsis, highly inclined elliptical lunar orbit.

Folta, David C.↗
NASA NTRS2001

Inviscid Flow Computations of the Orbital Sciences X-34 Over a Mach Number Range of 1.25 to 6.0

This report documents the results of an inviscid computational study conducted on the Orbital Sciences X-34 vehicle to compute its inviscid longitudinal aerodynamic characteristics over a Mach number range of 1.25 to 6.0. The unstructured grid software FELISA was used and th e aerodynamic characteristics were computed at Mach numbers 1.25, 1.6, 2.5, 4.0, 4.63, and 6.0, and an angle of attack range of -4 to 32 degrees. These results were compared with available aerodynamic data from wind tunnel test on X-34 models. The comparison showed excellent agreement in C(sub N). The computed pitching moment compared well at Mach numbers 2.5 and higher, and at angles of attack of up to 12 deg. The agreement was not good at higher angles of attack possibly due to viscous effects. At lower Mach numbers there were significant differences between computed and measured C(sub m) values. This could not be explained. Since the present computations are inviscid, the computed C(sub A) was consistently lower than the measured values as expected.

Prabhu, Ramadas K.↗
NASA NTRS2007

Report of the NASA Science Definition Team for the Mars Science Orbiter (MSO)

NASA is considering that its Mars Exploration Program (MEP) will launch an orbiter to Mars in the 2013 launch opportunity. To further explore this opportunity, NASA has formed a Science Definition Team (SDT) for this orbiter mission, provisionally called the Mars Science Orbiter (MSO). Membership and leadership of the SDT are given in Appendix 1. Dr. Michael D. Smith chaired the SDT. The purpose of the SDT was to define the: 1) Scientific objectives of an MSO mission to be launched to Mars no earlier than the 2013 launch opportunity, building on the findings for Plan A [Atmospheric Signatures and Near-Surface Change] of the Mars Exploration Program Analysis Group (MEPAG) Second Science Analysis Group (SAG-2); 2) Science requirements of instruments that are most likely to make high priority measurements from the MSO platform, giving due consideration to the likely mission, spacecraft and programmatic constraints. The possibilities and opportunities for international partners to provide the needed instrumentation should be considered; 3) Desired orbits and mission profile for optimal scientific return in support of the scientific objectives, and the likely practical capabilities and the potential constraints defined by the science requirements; and 4) Potential science synergies with, or support for, future missions, such as a Mars Sample Return. This shall include imaging for evaluation and certification of future landing sites. As a starting point, the SDT was charged to assume spacecraft capabilities similar to those of the Mars Reconnaissance Orbiter (MRO). The SDT was further charged to assume that MSO would be scoped to support telecommunications relay of data from, and commands to, landed assets, over a 10 Earth year period following orbit insertion. Missions supported by MSO may include planned international missions such as EXOMARS. The MSO SDT study was conducted during October - December 2007. The SDT was directed to complete its work by December 15, 2007. This rapid turn-around was required in order to allow time to prepare an Announcement of Opportunity (AO) for science investigations, to be released in early 2008.

Smith, Michael↗
NASA NTRS2008

The Mars Science Orbiter Concept

This paper describes a possible mission concept for the 2013 Mars launch opportunity. The actual NASA Mars Program plans and implementation may differ.

Advanced Mars Studies↗
NASA NTRSDate not supplied

How Can We Efficiently Build A Spacecraft That Has Longevity? Experiences From the GSFC Perspective to Inform Lunar Exploration and Science Orbiter’s Architecture

Recent successes in NASA planetary science missions has shown that long-lived missions can yield significant science return. These missions, despite their longevity, were not planned to operate beyond their deign life. For example, the Lunar Reconnaissance Orbiter has a design life of 2-3 years, and yet we are entering its 14th year on orbit. Spacecraft longevity in practice has been related to: 1) mission class; 2) did we test it long enough and resolve all the anomalies to be beyond the early failure curve; 3) consumables budgets, and 4) how components are de-signed and tested. Mission class has been perceived as one of the primary ways to drive longevity. But higher mission class is a significant cost and mission driver. Parts selection only from the limited military standard parts and extensive parts qualification adds to development time and cost. Largely redundant (often erroneously interpreted as fully redundant) adds to launch mass and increases testing complexity (which may reduce the amount of testing in the nominal con-figuration). Lower Risk Posture (reflected in a higher mission class) drives significant additional processes and quality assurance analyses. Higher mission class also drives significantly greater sparing and life testing costs. This discussion focuses on whether this is indeed the best way to achieve longevity efficiently or whether there are more efficient ways to achieve longevity. Empirical evidence shows that lower mission classes that implement effective risk reduction and selective redundancy generally results in long life at a lower Spacecraft bus cost. By evaluating redundancy careful-ly, spacecraft cost can be lowered which can enable a more capable payload. Risk of the mission is highly dependent on the complexity of the mission and is only loosely dependent on Class of Mission. High complexity missions can have many single point failures and require the development of new technology, while lower class mission can have lower risk by baselining or incorporating: (selective) redundancy, fault-tolerant design, design for minimum risk, ability to reset, and/or design for graceful degradation. The team met with Goddard Space Flight Center Space Systems Mission Operations leaders to discuss which avionics have proven in flight to be the most reliable and which have had lifetime issues. The paper proposes a list of components to focus on for redundancy for a long-lived lunar mission. Reliability numbers are presented for both single string and dual string missions. The history of life-time performance versus planned mission life according to mission class is presented. A mission with well thought out selective redundancy and effective on the ground test program, can be expected to last well beyond its mission lifetime and provided enhanced return for the community. This approach minimizes project expenditures that do not retire significant risk and allows the project to focus risk mitigation efforts on those risks that will have a significant likelihood of threatening mission success. This is aided by keeping the amount of technology maturation and mission complexity low, while focusing effort on ensuring all component stress-ing parameters well within the bounds of their capabilities.

Lessons Learned↗
NASA NTRS1972

Apollo lunar orbital science.

The Apollo 15 and 16 experiments are discussed. The formal scientific experiments conducted in lunar orbit are listed in an appendix. The gamma ray spectrometer performed a compositional survey of the upper 30 cm of the lunar surface by observing the gamma rays emitted during the radioactive decay of naturally occurring radioisotopes and their daughter products, and of the radioisotopes produced by cosmic ray bombardment of lunar surface materials. The results from the Apollo 15 X-ray spectrometer have shown excellent correlation between returned samples and gross lunar features. The various systems and experiments are described in detail.

Roberson, F. I.↗
NASA NTRS2012

Trajectory Options for a Potential Mars Mission Combining Orbiting Science, Relay and a Sample Return Rendezvous Demonstration

Mars sample return is a major scientific goal of the 2011 US National Research Council Decadal Survey for Planetary Science. Toward achievement of this goal, recent architecture studies have focused on several mission concept options for the 2018/2020 Mars launch opportunities. Mars orbiters play multiple roles in these architectures such as: relay, landing site identification/selection/certification, collection of on-going or new measurements to fill knowledge gaps, and in-orbit collection and transportation of samples from Mars to Earth. This paper reviews orbiter concepts that combine these roles and describes a novel family of relay orbits optimized for surface operations support. Additionally, these roles provide an intersection of objectives for long term NASA science, human exploration, technology development and international collaboration.

surface operations↗
NASA NTRS2021

Aerocapture Trajectories for Earth Orbit Technology Demonstration and Orbiter Science Missions at Venus, Earth, Mars, and Neptune

The use of aerocapture to provide the delta-V needed to capture a spacecraft into orbit can provide significant fuel savings compared to a propulsive orbit insertion. The aerocapture guidance method described in this paper uses drag modulation, which deploys a drag skirt in the atmosphere until the amount of delta-V needed to capture into the desired orbit is obtained, at which time the drag skirt is separated from the spacecraft. By modulating the time of drag skirt jettison, the vehicle can target a specific orbit apoapsis in the presence of uncertainties such as entry targeting errors and unknown atmospheric conditions.

Jadusingh, Matt↗