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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 577 records · Page 32
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NASA NTRS2007

Lessons Learned from Seven Space Shuttle Missions

Much can be learned from well-written descriptions of the technical and organizational factors that lead to an accident. Subsequent analysis by third parties of investigation reports and associated evidence collected during the investigations can lead to additional insight. Much can also be learned from documented close calls that do not result in loss of life or a spacecraft, such as the Mars Exploration Rover Spirit software anomaly, the SOHO mission interruption, and the NEAR burn anomaly. Seven space shuttle incidents fall into the latter category: Rendezvous Target Failure On STS-41B; Rendezvous Radar Anomaly and Trajectory Dispersion-STS-32 ;Rendezvous Lambert Targeting Anomaly-STS-49; Rendezvous Lambert Targeting Anomaly-STS-51; Zero Doppler Steering Maneuver Anomaly-STS-59; Excessive Propellant Consumption During Rendezvous-STS-69; Global Positioning System Receiver and Associated Shuttle Flight Software Anomalies-STS-91 Procedural work-arounds or software changes prevented them from threatening mission success. Extensive investigations, which included the independent recreation of the anomalies by multiple Shuttle Program organizations, were the key to determining the cause, accurately assessing risk, and identifying software and software process improvements. Lessons learned from these incidents not only validated long-standing operational best practices, but serve to promote discussion and mentoring among Program personnel and are applicable to future space flight programs.

Goodman, John↗
NASA NTRS1975

Analytical study of spacecraft deposition contamination by internal reflection spectroscopy

The in-situ vacuum chamber designed around the IRE was tested under vacuum. The contamination analyzer system was tested using two liquid contaminant materials. The analysis of the contaminants on the SL-4 rendezvous window shows that the material has methyl silicone, hydroxyl, and carbonyl radicals. The analysis of the outgassing material from Shuttle TPS shows silicone as the primary product.

Mookherji, T. K.↗
NASA NTRS1990

Upcoming planetary missions and the applicability of high temperature superconductor bolometers

Planetary missions to Mars and beyond can last 11 years and longer, making impractical the use of stored cryogens. Passive radiative coolers and single-stage mechanical coolers remain possibilities. CRAF and CASSINI, both using the newly developed Mariner Mark 2 spacecraft, will be the next outer planet missions after Galileo; they are intended to provide information on the origin and evolution of the solar system. CRAF is a cometary rendezvous mission slated for a 1994 launch. CASSINI has been chosen by ESA and will be launched by a Titan 4/Centaur in 1996. It will fly by Jupiter in 2000, inject an ESA-supplied probe into Titan in 2002, and take data in Saturn orbit from 2002 to 2006. NASA/Goddard is currently developing a prototype Fourier transform spectrometer (CIRS) that will be proposed for the CASSINI mission. The baseline infrared detectors for CIRS are HgCdTe to 16 microns and Schwarz-type thermopiles from 16 to 1000 microns. The far infrared focal plane could be switched from thermopiles to high temperature superconductor (HTS) bolometers between now and 1996. An HTS bolometer could be built using the kinetic inductance effect, or the sharp resistance change at the transition. The transition-edge bolometer is more straightforward to implement and initial efforts at NASA/Goddard are directed to that device. A working device was made and tested in early 1989. It also has somewhat elevated noise levels below 100 Hz. Upcoming efforts will center on reducing the time constant of the HTS bolometer by attempting to deposit an HTS film on a diamond substrate, and by thinning SrTiO3 substrates. Attempts will be made to improve the film quality to reduce th 1/f noise level, and to improve the thermal isolation to increase the bolometer sensitivity. An attempt is being made to deposit good-quality HTS films on diamond films using an MOCVD technique.

Brasunas, J.↗
NASA NTRS1995

STS-84 Day 08 Highlights

On this eighth day of the STS-84 mission, the flight crew, Cmdr. Charles J. Precourt, Pilot Eileen M. Collins, Payload Cmdr, Jean-Francois Clervoy (ESA), Mission Specialists Edward T. Lu, Carlos I. Noriega, Elena V. Kondakova, Jerry M. Linenger (download), and C. Michael Foale (upload) sing 'The Cosmonauts' Song' to Mir-23 crew members Vasily Tsibliev, Alexander Lazutkin and astronaut Mike Foale, who is beginning his four-month research mission on Mir. Foale and his new crewmates played music as Atlantis departed following the joint phase of the flight. Atlantis' undocking from Mir was modified from previous joint missions in that a flyaround of the station for photographic purposes was not conducted. Instead, Pilot Eileen Collins guided Atlantis below the Mir after the two spacecraft completed their physical separation, stopping three times at distances of 90, 300 and 1,500 feet to collect data from a European sensor device designed to assist future rendezvous of a proposed European Space Agency resupply vehicle with the International Space Station. Once the data collection was completed, the shuttle took advantage of natural orbital mechanics to drift beneath and out in front of Mir.

Source record↗
NASA NTRS2004

Mission Steering Profiles of Outer Planetary Orbiters Using Radioisotope Electric Propulsion

Radioisotope Electric Propulsion (REP) has the potential to enable small spacecraft to orbit outer planetary targets with trip times comparable to flyby missions. The ability to transition from a flyby to an orbiter mission lies in the availability of continuous low power electric propulsion along the entire trajectory. The electric propulsion system s role is to add and remove energy from the spacecraft s trajectory to bring it in and out of a heliocentric hyperbolic escape trajectory for the outermost target bodies. Energy is added and the trajectory is reshaped to rendezvous with the closer-in target bodies. Sample REP trajectories will be presented for missions ranging for distances from Jupiter orbit to the Pluto-Kuiper Belt.

Fiehler, Douglas↗
NASA NTRS2013

An Electric Propulsion "Shepherd" for Active Debris Removal that Utilizes Ambient Gas as Propellant

There is a growing consensus among the space debris technical community that limiting the long ]term growth of debris in Low-Earth Orbit (LEO) requires that space users limit the accumulation of mass in orbit. This is partially accomplished by mitigation measures for current and future LEO systems, but there is now interest in removing mass that has already accumulated in LEO from more than 50 years of space activity (termed "Active Debris Removal", or ADR). Many ADR proposals face complex technical issues of how to grapple with uncooperative targets. Some researchers have suggested the use of conventional ion thrusters to gently "blow" on objects to gradually change their orbits, without ever having to come into physical contact with the target. The chief drawback with these methods is the cost per object removed. Typically, a space "tug" or an ion-drive "shepherd" can only remove a few objects per mission due to limited propellant. Unless a costeffective way that removes tens of objects per mission can be found, it is not clear that any of the ideas so far proposed will be economically viable. In this paper, a modified version of the ion-drive "shepherd" is proposed that uses ambient atmospheric gases in LEO as propellant for the ion drives. This method has the potential to greatly extend the operational lifetime of an ADR mission, as the only mission limit is the lifetime of the components of the satellite itself, not on its fuel supply. An ambient-gas ion-drive "shepherd" would the local atmospheric drag on an object by ionizing and accelerating the ambient gas the target would have encountered anyway, thereby hastening its decay. Also, the "shepherd" satellite itself has a great deal of flexibility to maneuver back to high altitude and rendezvous with its next target using the ion drive not limited by fuel supply. However, the amount of available ambient gas is closely tied to the altitude of the spacecraft. It may be possible to use a "hybrid" approach that supplements high-altitude ion-drive operations with stored gas, and transitions to ambient gas at lower altitudes. This paper will include realistic numbers on the estimated times needed to deorbit objects from different orbit regimes using drives that either partially or completely take advantage of ambient gas. It will conclude with recommendations on whether this is a viable candidate for future ADR efforts.

Matney, Mark J.↗
NASA NTRS2013

An Electric Propulsion "Shepherd" for Active Debris Removal that Utilizes Ambient Gas as Propellant

There is a growing consensus among the space debris technical community that limiting the long-term growth of debris in Low-Earth Orbit (LEO) requires that space users limit the accumulation of mass in orbit. This is partially accomplished by mitigation measures for current and future LEO systems, but there is now interest in removing mass that has already accumulated in LEO from more than 50 years of space activity (termed "Active Debris Removal", or ADR). Many ADR proposals face complex technical issues of how to grapple with uncooperative targets. Some researchers have suggested the use of conventional ion thrusters to gently "blow" on objects to gradually change their orbits, without ever having to come into physical contact with the target. The chief drawback with these methods is the cost per object removed. Typically, a space "tug" or an ion-drive "shepherd" can only remove a few objects per mission due to limited propellant. Unless a cost-effective way that removes tens of objects per mission can be found, it is not clear that any of the ideas so far proposed will be economically viable. In this paper, a modified version of the ion-drive "shepherd" is proposed that uses ambient atmospheric gases in LEO as propellant for the ion drives. This method has the potential to greatly extend the operational lifetime of an ADR mission, as the only mission limit is the lifetime of the components of the satellite itself, not on its fuel supply. An ambient-gas ion-drive "shepherd" would enhance the local atmospheric drag on an object by ionizing and accelerating the ambient gas the target would have encountered anyway, thereby hastening its decay. Also, the "shepherd" satellite itself has a great deal of flexibility to maneuver back to high altitude and rendezvous with its next target using the ion drive not limited by fuel supply. However, the amount of available ambient gas is closely tied to the altitude of the spacecraft. It may be possible to use a "hybrid" approach that supplements high-altitude ion-drive operations with stored gas, and transitions to ambient gas at lower altitudes. This paper will include realistic numbers on the estimated times needed to deorbit objects from different orbit regimes using drives that either partially or completely take advantage of ambient gas. It will conclude with recommendations on whether this is a viable candidate for future ADR efforts.

Matney, Mark↗
NASA NTRS2015

In-Flight Operation of the Dawn Ion Propulsion System Through Survey Science Orbit at Ceres

The Dawn mission, part of NASA's Discovery Program, has as its goal the scientific exploration of the two most massive main-belt objects, Vesta and Ceres. The Dawn spacecraft was launched from the Cape Canaveral Air Force Station on September 27, 2007 on a Delta-II 7925H- 9.5 (Delta-II Heavy) rocket that placed the 1218-kg spacecraft onto an Earth-escape trajectory. On-board the spacecraft is an ion propulsion system (IPS) developed at the Jet Propulsion Laboratory which will provide a total delta V of 11 km/s for the heliocentric transfer to Vesta, orbit capture at Vesta, transfer between Vesta science orbits, departure and escape from Vesta, heliocentric transfer to Ceres, orbit capture at Ceres, and transfer between Ceres science orbits. Full-power thrusting from December 2007 through October 2008 was used to successfully target a Mars gravity assist flyby in February 2009 that provided an additional delta V of 2.6 km/s. Deterministic thrusting for the heliocentric transfer to Vesta resumed in June 2009 and concluded with orbit capture at Vesta on July 16, 2011. From July 2011 through September 2012 the IPS was used to transfer to all the different science orbits at Vesta and to escape from Vesta orbit. Cruise for a rendezvous with Ceres began in September 2012 and concluded with the start of the approach to Ceres phase on December 26, 2015, leading to orbit capture on March 6, 2015. Deterministic thrusting continued during approach to place the spacecraft in its first science orbit, called RC3, which was achieved on April 23, 2015. Following science operations at RC3 ion thrusting was resumed for twenty-five days leading to arrival to the next science orbit, called survey orbit, on June 3, 2015. The IPS will be used for all subsequent orbit transfers and trajectory correction maneuvers until completion of the primary mission in approximately June 2016. To date the IPS has been operated for over 46,774 hours, consumed approximately 393 kg of xenon, and provided a delta V of over 10.8 km/s to the spacecraft. The IPS performance characteristics are very close to the expected performance based on analysis and testing performed pre-launch. This paper provides an overview of Dawn's mission objectives and the results of Dawn IPS mission operations through arrival at the second science orbit at Ceres.

Garner, Charles E.↗
NASA NTRS1971

Guidance and control of flight vehicles

Progress reports on guidance and attitude control mechanisms of different flight vehicles are presented. The vehicles considered include orbiting spacecraft, supersonic aircraft, and general aviation aircraft. Data also cover orbital transfer using low thrust, automatic landing logic for aircraft, optimal and three dimensional turns for supersonic aircraft, and orbital rendezvous.

Breakwell, J. V.↗
NASA NTRS2001

Apollo 13 Facts

Footage is seen of the Earth from the Apollo 13 spacecraft as it travels towards the Moon. The crew, James Lovell, Jr., John Swigert, Jr., and Fred Haise, Jr., are shown performing various on-orbit activities. The Lunar Module rendezvous and docking, tunnel repressurization, and S4-B separation are also seen.

Source record↗
NASA NTRSDate not supplied

Linear Covariance Techniques to Analyze a Multi-Vehicle, Multi-GN&C System with Applications to Rendezvous in a Near Rectilinear Halo Orbit

Typically for a rendezvous and docking scenario, only a single vehicle is considered the active vehicle. As a result, the target vehicle’s GN&C system is typically not a factor in the integrated performance analysis as it assumes a passive role. However, for upcoming NASA Artemis missions the active vehicle is both the chaser and target spacecraft. In addition, both vehicles are cooperating with one another, sharing telemetry data, and relying on data extracted from the GN&C system of the other. The process also includes a team of ground support personnel in mission control are tracking and monitoring each spacecraft uplinking state estimates and targeting solutions to support mission operations and enhance the onboard flight system performance. Being able to quickly analyze the impact of both vehicles with two different GN&C systems that also interact with the ground that does its own navigation and targeting uploads is critical. This paper outlines how to perform this rapid analysis using linear covariance techniques and applies them to a rendezvous scenario initiated in low lunar orbit and completed in a Near Rectilinear Halo Orbit (NRHO) representative of the NASA Artemis III mission.

Linear Covariance Analysis↗
NASA NTRS1999

On Orbit Testing of the Video Guidance Sensor

The Video Guidance Sensor (VGS), part of NASA's Automated Rendezvous and Capture program, was flown on Shuttle mission STS-95 in October of 1998 to orbitally test the functional characteristics of the VGS. This was the second flight of the VGS (the first flight was in 1997 on STS-87), and this time long-range tracking data was gathered during the experiment. The flight experiment sensor was designed to operate from 1.5 meter range out to 110 meter range, with a field-of-view of 16 X 21 degrees. The VGS tracked its target at a 5 Hz rate and returned 6-degree-of-freedom information on the target's position and attitude relative to the sensor. The VGS was mounted in the Shuttle cargo bay, and its target was mounted on the Spartan spacecraft being carried on this mission. The orbital testing of the VGS included operations with the target on the Shuttle's Remote Manipulator System (RMS) at the start of the 10-day mission, long-range data during the Shuttle rendezvous with the Spartan two days later, and some more RMS operations later in the mission. The data returned from the orbital testing included VGS diagnostics, acquisition, and tracking data, RMS positions, hand-held laser range data, tapes of the data from the VGS video camera, and orbital positioning data from the Spartan and the Shuttle to allow correlation of the VGS data with orbital best-estimate-of-truth data. The Video Guidance Sensor performed well in all phases of the testing, and the VGS is being incorporated into the ground testing of a complete automated rendezvous and docking system. Work on the development of the next generation VGS is continuing

Howard, Richard T.↗
NASA NTRS1986

Approach strategy for the rendezvous with comet Tempel 2

The methods used to locate the comet nucleus and safety navigate through the debris surrounding the comet are discussed. The approach to the nucleus is in two parts: a high-speed (1.9 km/s) leg which ends with the spacecraft 200,000 km sunward of the nucleus, and a low-speed (25 m/s) leg which traverses the 200,000 km to the nucleus. The nucleus will be recovered by the narrow-angle camera 30 days prior to the completion of the high-speed inbound leg. The debris density is low enough to permit the safe passage of the spacecraft through the surrounding cometary debris.

Kakuda, R. Y.↗
NASA NTRS2013

Mars Sample Return Using Commercial Capabilities: Mission Architecture Overview

Mars Sample Return (MSR) is the highest priority science mission for the next decade as recommended by the recent Decadal Survey of Planetary Science. This paper presents an overview of a feasibility study for a MSR mission. The objective of the study was to determine whether emerging commercial capabilities can be used to reduce the number of mission systems and launches required to return the samples, with the goal of reducing mission cost. The major element required for the MSR mission are described and include an integration of the emerging commercial capabilities with small spacecraft design techniques; new utilizations of traditional aerospace technologies; and recent technological developments. We report the feasibility of a complete and closed MSR mission design using the following scenario that covers three synodic launch opportunities, beginning with the 2022 opportunity: A Falcon Heavy injects a SpaceX Red Dragon capsule and trunk onto a Trans Mars Injection (TMI) trajectory. The capsule is modified to carry all the hardware needed to return samples collected on Mars including a Mars Ascent Vehicle (MAV); an Earth Return Vehicle (ERV); and hardware to transfer a sample collected in a previously landed rover mission to the ERV. The Red Dragon descends to land on the surface of Mars using Supersonic Retro Propulsion (SRP). After previously collected samples are transferred to the ERV, the single-stage MAV launches the ERV from the surface of Mars to a Mars phasing orbit. The MAV uses a storable liquid, pump fed bi-propellant propulsion system. After a brief phasing period, the ERV, which also uses a storable bi-propellant system, performs a Trans Earth Injection (TEI) burn. Once near Earth the ERV performs Earth and lunar swing-bys and is placed into a Lunar Trailing Orbit (LTO0 - an Earth orbit, at lunar distance. A later mission, using a Dragon and launched by a Falcon Heavy, performs a rendezvous with the ERV in the lunar trailing orbit, retrieves the sample container and breaks the chain of contact with Mars by transferring the sample into a sterile and secure container. With the sample contained, the retrieving spacecraft, makes a controlled Earth re-entry preventing any unintended release of pristine Martian materials into the Earth's biosphere. Other capsule type vehicles and associated launchers may be applicable. The analysis methods employed standard and specialized aerospace engineering tools. Mission system elements were analyzed with either direct techniques or by using parametric mass estimating relationships (MERs). The architecture was iterated until overall mission convergence was achieved on at least one path. Subsystems analyzed in this study include support structures, power system, nose fairing, thermal insulation, actuation devices, MAV exhaust venting, and GN&C. Best practice application of loads, mass growth contingencies, and resource margins were used. For Falcon Heavy capabilities and Dragon subsystems we utilized publically available data from SpaceX; published analyses from other sources; as well as our own engineering and aerodynamic estimates. Earth Launch mass is under 11 mt, which is within the estimated capability of a Falcon Heavy, with margin. Total entry masses between 7 and 10 mt were considered with closure occurring between 9 and 10 mt. Propellant mass fractions for each major phase of the EDL - Entry, Terminal Descent, and Hazard Avoidance - have been derived. An assessment of the entry conditions on the thermal protection system (TPS), currently in use for Dragon missions, has been made. And shows no significant stressors. A useful mass of 2.0 mt is provided and includes mass growth allowances for the MAV, the ERV, and mission unique equipment. We also report on alternate propellant options for the MAV and options for the ERV, including propulsion systems; crewed versus robotic retrieval mission; as well as direct Earth entry. International Planetary Protection Policies as well as verifiable means of compliance will have a large impact on any MSR mission design. We identify areas within our architecture where such impacts occur. This work shows that emerging commercial capabilities can be used to effectively integrated into a mission to achieve an important planetary science objective.

Mars Sample Return↗
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, Stephen D.↗