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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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At least 271 records · Page 15
All sourcesNASA NTRSDOE OSTI
NASA NTRS1995

Guidance, navigation, and control study for a solar electric propulsion spacecraft

A preliminary investigation of a lunar-comet rendezvous mission using a solar electric propulsion (SEP) spacecraft was performed in two phases.The first phase involved exploration of the moon and the second involved rendezvous with a comet. The initial phase began with a chemical propulsion translunar injection and chemical insertion into a lunar orbit, followed by a low thrust SEP transfer to a circular, polar, low-lunar orbit. After collecting scientific data at the moon, the SEP spacecraft performed a spiral lunar escape maneuver to begin the interplanetary leg of the mission. After escape from the Earth-moon system, the SEP spacecraft maneuvered in interplanetary space and performed a rendezvous with a comet.The immediate goal of this study was to demonstrate the feasibility of using a low-thrust SEP spacecraft for orbit transfer to both the moon and a comet. Another primary goal was to develop a computer optimization code which would be robust enough to obtain minimum-fuel rendezvous trajectories for a wide range of comets.

Kluever, Craig A.↗
NASA NTRS1995

Optimal lunar trajectories for a combined chemical-electric propulsion spacecraft

Spacecraft which utilize electric propulsion (EP) systems are capable of delivering a greater payload fraction compared to spacecraft using conventional chemical propulsion systems. Several researchers have investigated numerous applications of low-thrust EP including a manned Mars mission, scientific missions to the outer planets, and lunar missions. In contrast, the study of optimal combined high and low-thrust spacecraft trajectories has been limited. In response to the release of NASA's 1994 Announcement of Opportunity (AO) for Discovery class interplanetary exploration missions, a preliminary investigation of a lunar comet rendezvous mission using a solar electric propulsion (SEP) spacecraft was performed. The Discovery mission (eventually named Diana) was envisioned to be a two-phase scientific exploration mission: the first phase involved exploration of the moon and second phase involved rendezvous with a comet. The initial phase began with a chemical propulsion translunar injection and chemical insertion into a lunar orbit, followed by a low-thrust SEP transfer to a circular, polar, low-lunar orbit (LLO). After scientific data was collected at the moon, the SEP spacecraft performed a spiral lunar escape maneuver to begin the interplanetary leg of the mission. After escape from the Earth-moon system, the SEP spacecraft maneuvered in interplanetary space and performed a rendezvous with a short period comet. An initial study that demonstrated the feasibility of using EP for the lunar and comet orbit transfer was performed under the grant NAG3-1581. This final report is a continuation of the initial research efforts in support of the Discovery mission proposal that was submitted to NASA Headquarters in October 1994. Section 2 discusses the lunar orbit transfer phase of the Diana mission which involves both chemical and electric propulsion stages. Section 3 discusses the chemical lunar orbit insertion (LOI) burn optimization. Finally, section 4 presents the conclusions of this research effort.

Kluever, Craig A.↗
NASA NTRS2006

Liquid Hydrogen Sensor Considerations for Space Exploration

The on-orbit management of liquid hydrogen planned for the return to the moon will introduce new considerations not encountered in previous missions. This paper identifies critical liquid hydrogen sensing needs from the perspective of reliable on-orbit cryogenic fluid management, and contrasts the fundamental differences in fluid and thermodynamic behavior for ground-based versus on-orbit conditions. Opportunities for advanced sensor development and implementation are explored in the context of critical Exploration Architecture operations such as on-orbit storage, docking, and trans-lunar injection burn. Key sensing needs relative to these operations are also examined, including: liquid/vapor detection, thermodynamic condition monitoring, mass gauging, and leak detection. Finally, operational aspects of an integrated system health management approach are discussed to highlight the potential impact on mission success.

Moran, Matthew E.↗
NASA NTRS2004

Systems engingeering for the Kepler Mission : a search for terrestrial planets

The Kepler mission will launch in 2007 and determine the distribution of earth-size planets (0.5 to 10 earth masses) in the habitable zones (HZs) of solar-like stars. The mission will monitor > 100,000 dwarf stars simultaneously for at least 4 years. Precision differential photometry will be used to detect the periodic signals of transiting planets. Kepler will also support asteroseismology by measuring the pressure-mode (p-mode) oscillations of selected stars. Key mission elements include a spacecraft bus and 0.95 meter, wide-field, CCD-based photometer injected into an earth-trailing heliocentric orbit by a 3-stage Delta II launch vehicle as well as a distributed Ground Segment and Follow-up Observing Program. The project is currently preparing for Preliminary Design Review (October 2004) and is proceeding with detailed design and procurement of long-lead components. In order to meet the unprecedented photometric precision requirement and to ensure a statistically significant result, the Kepler mission involves technical challenges in the areas of photometric noise and systematic error reduction, stability, and false-positive rejection. Programmatic and logistical challenges include the collaborative design, modeling, integration, test, and operation of a geographically and functionally distributed project. A very rigorous systems engineering program has evolved to address these challenges. This paper provides an overview of the Kepler systems engineering program, including some examples of our processes and techniques in areas such as requirements synthesis, validation & verification, system robustness design, and end-to-end performance modeling.

systems engineering↗
NASA NTRS2014

A High Power Solar Electric Propulsion - Chemical Mission for Human Exploration of Mars

Recently Solar Electric Propulsion (SEP) as a main propulsion system has been investigated as an option to support manned space missions to near-Earth destinations for the NASA Gateway spacecraft. High efficiency SEP systems are able to reduce the amount of propellant long duration chemical missions require, ultimately reducing the required mass delivered to Low Earth Orbit (LEO) by a launch vehicle. However, for long duration interplanetary Mars missions, using SEP as the sole propulsion source alone may not be feasible due to the long trip times to reach and insert into the destination orbit. By combining an SEP propulsion system with a chemical propulsion system the mission is able to utilize the high-efficiency SEP for sustained vehicle acceleration and deceleration in heliocentric space and the chemical system for orbit insertion maneuvers and trans-earth injection, eliminating the need for long duration spirals. By capturing chemically instead of with low-thrust SEP, Mars stay time increases by nearly 200 days. Additionally, the size the of chemical propulsion system can be significantly reduced from that of a standard Mars mission because the SEP system greatly decreases the Mars arrival and departure hyperbolic excess velocities (V(sub infinity)).

Solar Electric Propulsion↗
NASA NTRSDate not supplied

Optimization of the Lunar Icecube Trajectory Using Stochastic Global Search and Multi-Point Shooting

Lunar IceCube is a 6U cubesat that will launch on NASA’s Artemis 1 mission in 2021. Lunar IceCube will separate from Artemis 1 shortly after trans-lunar injection (TLI) and travel to its science orbit about the moon using its Busek Ion Thruster 3 (BIT-3) propulsion system. This paper describes a technique to rapidly design Lunar IceCube trajectories using the monotonic basin hopping (MBH) stochastic global search algorithm, along with low- and high-fidelity multi-point shooting transcriptions. This technique allows the Lunar IceCube team to rapidly adapt to changing initial conditions, spacecraft properties, and operational constraints.

optimization↗
NASA NTRS1996

Substorm associated ion acceleration in the dynamic magnetotail

Ion acceleration and flux increase associated with substorm energetic particle injections are investigated on the basis of geosynchronous observations and test proton orbits in the dynamic fields of a three dimensional magnetohydrodynamic simulation of neutral line formation and dipolarization in the magnetotail. The energetic particle flux changes obtained from the test particle orbits agree well with observations that demonstrate rapid ion flux increases at energies of above 20 keV. The injection region inferred from the test particles has a sharp earthward boundary and a sharp ragged tailward boundary. The earthward portion of the enhanced ion flux can be traced to the enhanced cross-tail electric field associated with the near-earth x-type neutral line. Due to the rapid earthward motion of accelerated ions away from the neutral line, this boundary is displaced earthward to where the energetic ions become more adiabatic in the stronger dipolar field.

Birn, J.↗
NASA NTRS2008

Methods of Helium Injection and Removal for Heat Transfer Augmentation

While augmentation of heat transfer from a test article by helium gas at low pressures is well known, the method is rarely employed during space simulation testing because the test objectives are to simulate an orbital thermal environment. Test objectives of cryogenic optical testing at Marshall Space Flight Center's X-ray Calibration Facility (XRCF) have typically not been constrained by orbital environment parameters. As a result, several methods of helium injection have been utilized at the XRCF since 1999 to decrease thermal transition times. A brief synopsis of these injection (and removal) methods including will be presented.

Kegley, Jeffrey↗
NASA NTRSDate not supplied

Composition and Climate Impacts of Increasing Launches to Low Earth Orbit

We present simulations of an Earth system model in which launch vehicles inject gaseous and particulate combustion products into the stratosphere. We considered two plausible scenarios representative for the year 2050, one with 1000 launches per year and another with 10 times more, all of which assuming heavy lift (80 tons to low Earth orbit; LEO) launches using methane as a proxy for liquefied natural gas (LNG) fuel. An industry-standard plume flowfield model was used to predict emission of gaseous species including carbon monoxide, nitrogen oxides, and water vapor. Black carbon was introduced assuming vacuum emission equal 25% of equivalent kerosene engine. The gas emissions showed marginal or insignificant changes in atmospheric composition, while the black carbon emission was found to be the most influential component. Feedbacks involve tropopause and stratospheric warming, followed by a moistening of the stratosphere via tropospheric water vapor intrusions, together with a reduction in global albedo. The fact that LNG-fueled rockets could be a concern from aerosol emissions in our simulations, despite burning more efficiently than conventional fuels, underscores the need to better understand actual black carbon emissions from rocket engines in order to accurately predict the global impacts of launch vehicles on future climate.

Earth system model↗
NASA NTRS2008

Methods of Helium Injection and Removal for Heat Transfer Augmentation

While augmentation of heat transfer from a test article by helium gas at low pressures is well known, the method is rarely employed during space simulation testing because the test objectives usually involve simulation of an orbital thermal environment. Test objectives of cryogenic optical testing at Marshall Space Flight Center's X-ray Cryogenic Facility (XRCF) have typically not been constrained by orbital environment parameters. As a result, several methods of helium injection have been utilized at the XRCF since 1999 to decrease thermal transition times. A brief synopsis of these injection (and removal) methods including will be presented.

Haight, Harlan↗
NASA NTRS1981

Space shuttle launch era spacecraft injection errors and DSN initial acquisition

The initial acquisition of a spacecraft by the Deep Space Network (DSN) is a critical mission event. This results from the importance of rapidly evaluating the health and trajectory of a spacecraft in the event that immediate corrective action might be required. Further, the DSN initial acquisition is always complicated by the most extreme tracking rates of the mission. The DSN initial acquisition characteristics will change considerably in the upcoming space shuttle launch era. How given injection errors at spacecraft separation from the upper stage launch vehicle (carried into orbit by the space shuttle) impact the DSN initial acquisition, and how this information can be factored into injection accuracy requirements to be levied on the Space Transportation System (STS) is addressed. The approach developed begins with the DSN initial acquisition parameters, generates a covariance matrix, and maps this covariance matrix backward to the spacecraft injection, thereby greatly simplifying the task of levying accuracy requirements on the STS, by providing such requirements in a format both familiar and convenient to STS.

Khatib, A. R.↗
NASA NTRS2007

First In-Orbit Experience of TerraSAR-X Flight Dynamics Operations

TerraSAR-X is an advanced synthetic aperture radar satellite system for scientific and commercial applications that is realized in a public-private partnership between the German Aerospace Center (DLR) and the Astrium GmbH. TerraSAR-X was launched at June 15, 2007 on top of a Russian DNEPR-1 rocket into a 514 km sun-synchronous dusk-dawn orbit with an 11-day repeat cycle and will be operated for a period of at least 5 years during which it will provide high resolution SAR-data in the X-band. Due to the objectives of the interferometric campaigns the satellite has to comply to tight orbit control requirements, which are formulated in the form of a 250 m toroidal tube around a pre-flight determined reference trajectory (see [1] for details). The acquisition of the reference orbit was one of the main and key activities during the Launch and Early Orbit Phase (LEOP) and had to compensate for both injection errors and spacecraft safe mode attitude control thruster activities. The paper summarizes the activities of GSOC flight dynamics team during both LEOP and early Commissioning Phase, where the main tasks have been 1) the first-acquisition support via angle-tracking and GPS-based orbit determination, 2) maneuver planning for target orbit acquisition and maintenance, and 3) precise orbit and attitude determination for SAR processing support. Furthermore, a presentation on the achieved results and encountered problems will be addressed.

Kahle, R.↗
NASA NTRS1986

System engineering study of electrodynamic tether as a spaceborne generator and radiator of electromagnetic waves in the ULF/ELF frequency band

The transmission and generation by orbiting tethered satellite systems of information carrying electromagnetic waves in the ULF/ELF frequency band to the Earth at suitably high signal intensities was examined and the system maintaining these intensities in their orbits for long periods of time without excessive onboard power requirements was investigated. The injection quantity power into electromagnetic waves as a function of system parameters such as tether length and orbital height was estimated. The basic equations needed to evaluate alternataing current tethered systems for external energy requirements are presented. The energy equations to tethered systems with various lengths, tether resistances, and radiation resistances, operating at different current values are applied. Radiation resistance as a function of tether length and orbital height is discussed. It is found that ULF/ELF continuously radiating systems could be maintained in orbit with moderate power requirements. The effect of tether length on the power going into electromagnetic waves and whether a single or dual tether system is preferable for the self-driven mode is discussed. It is concluded that the single tether system is preferable over the dual system.

Estes, R. D.↗
NASA NTRS1973

Differential transformation for satellite injection.

This paper provides an analytical method for computing the required orbital inclination change when the nominal orbital node rate is slightly modified to meet satellite mission requirements. A first-order theory retaining up to J sub 2 terms of the earth-oblateness perturbation has been found adequate to generate a four-decimal digit accuracy for required orbital changes. A numerical example of an actual ITOS series prelaunch analysis has been given to illustrate computational steps.

Sugai, I.↗
NASA NTRS1988

LLOFX earth orbit to lunar orbit delta V estimation program user and technical documentation

The LLOFX computer program calculates in-plane trajectories from an Earth-orbiting space station to Lunar orbit in such a way that the journey requires only two delta V burns (one to leave Earth circular orbit and one to circularize into Lunar orbit). The program requires the user to supply the Space Station altitude and Lunar orbit altitude (in km above the surface), and the desired time of flight for the transfer (in hours). It then determines and displays the trans-Lunar injection (TLI) delta V required to achieve the transfer, the Lunar orbit insertion (LOI) delta V required to circularize the orbit around the Moon, the actual time of flight, and whether the transfer orbit is elliptical or hyperbolic. Return information is also displayed. Finally, a plot of the transfer orbit is displayed.

Source record↗
NASA NTRS2009

NASA Constellation Program (CxP) Key Driving Requirements and Element Descriptions for International Architecture Working Group (IAWG) Functional Teams Human Transportation Cargo Transportation

The NASA Constellation uncrewed cargo mission delivers cargo to any designated location on the lunar surface (or other staging point) in a single mission. This capability is used to deliver surface infrastructure needed for lunar outpost construction, to provide periodic logistics resupply to support a continuous human lunar presence, and potentially deliver other assets to various locations.In the nominal mission mode, the Altair lunar lander is launched on Ares V into Low Earth Orbit (LEO), following a short Low Earth Orbit (LEO) loiter period, the Earth Departure Stage (EDS) performs the Trans Lunar Injection (TLI) burn and is then jettisoned. The Altair performs translunar trajectory correction maneuvers as necessary and performs the Lunar Orbit Insertion (LOI) burn. Altair then descends to the surface to land near a designated target, presumably in proximity to an Outpost location or another site of interest for exploration.Alternatively, the EDS and Altair Descent Stage could deliver assets to various staging points within their propulsive capabilities.

Martinez, Roland M.↗
NASA NTRS1978

Alterations in erythrocyte survival parameters in rats after 19.5 days aboard Cosmos 782

Rats were subjected to 19.5 days of weightless space flight aboard the Soviet biosatellite, Cosmos 782. Based on the output of CO-14, survival parameters of a cohort of erythrocytes labeled 15.5 days preflight were evaluated upon return from orbit. These were compared to vivarium control rats injected at the same time. Statistical evaluation indicates that all survival factors were altered by the space flight. The mean potential lifespan, which was 63.0 days in the control rats, was decreased to 59.0 days in the flight rats, and random hemolysis was increased three-fold in the flight rats. The measured size of the cohort was decreased, lending further support to the idea that hemolysis was accelerated during some portion of the flight. A number of factors that might be contributory to these changes are discussed, including forces associated with launch and reentry, atmospheric and environmental parameters, dietary factors, radiation, and weightlessness.

Leon, H. A.↗
NASA NTRS2012

Launch and Assembly Reliability Analysis for Human Space Exploration Missions

NASA's future human space exploration strategy includes single and multi-launch missions to various destinations including cis-lunar space, near Earth objects such as asteroids, and ultimately Mars. Each campaign is being defined by Design Reference Missions (DRMs). Many of these missions are complex, requiring multiple launches and assembly of vehicles in orbit. Certain missions also have constrained departure windows to the destination. These factors raise concerns regarding the reliability of launching and assembling all required elements in time to support planned departure. This paper describes an integrated methodology for analyzing launch and assembly reliability in any single DRM or set of DRMs starting with flight hardware manufacturing and ending with final departure to the destination. A discrete event simulation is built for each DRM that includes the pertinent risk factors including, but not limited to: manufacturing completion; ground transportation; ground processing; launch countdown; ascent; rendezvous and docking, assembly, and orbital operations leading up to trans-destination-injection. Each reliability factor can be selectively activated or deactivated so that the most critical risk factors can be identified. This enables NASA to prioritize mitigation actions so as to improve mission success.

Cates, Grant↗