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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 NTRS2016

An Efficient Approach for Mars Sample Return Using Emerging Commercial Capabilities

Mars Sample Return is the highest priority science mission for the next decade as recommended by the 2011 Decadal Survey of Planetary Science. This article presents the results of a feasibility study for a Mars Sample Return mission that efficiently uses emerging commercial capabilities expected to be available in the near future. The motivation of our study was the recognition that emerging commercial capabilities might be used to perform Mars Sample Return with an Earth-direct architecture, and that this may offer a desirable simpler and lower cost approach. The objective of the study was to determine whether these capabilities can be used to optimize the number of mission systems and launches required to return the samples, with the goal of achieving the desired simplicity. All of the major element required for the Mars Sample Return mission are described. Mission system elements were analyzed with either direct techniques or by using parametric mass estimating relationships. The analysis shows the feasibility of a complete and closed Mars Sample Return mission design based on the following scenario: A SpaceX Falcon Heavy launch vehicle places a modified version of a SpaceX Dragon capsule, referred to as "Red Dragon", onto a Trans Mars Injection trajectory. The capsule carries all the hardware needed to return to Earth Orbit samples collected by a prior mission, such as the planned NASA Mars 2020 sample collection rover. The payload includes a fully fueled Mars Ascent Vehicle; a fueled Earth Return Vehicle, support equipment, and a mechanism to transfer samples from the sample cache system onboard the rover to the Earth Return Vehicle. The Red Dragon descends to land on the surface of Mars using Supersonic Retropropulsion. After collected samples are transferred to the Earth Return Vehicle, the single-stage Mars Ascent Vehicle launches the Earth Return Vehicle from the surface of Mars to a Mars phasing orbit. After a brief phasing period, the Earth Return Vehicle performs a Trans Earth Injection burn. Once near Earth, the Earth Return Vehicle performs Earth and lunar swing-bys and is placed into a Lunar Trailing Orbit - an Earth orbit, at lunar distance. A retrieval mission then performs a rendezvous with the Earth Return Vehicle, 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 Martian materials into the Earth's biosphere. The mission can start in any one of three Earth to Mars launch opportunities, beginning in 2022.

Commercial↗
NASA NTRS1985

Suprathermal plasma observed on STS-3 Mission by plasma diagnostics package

Artificially produced electron beams were used extensively during the past decade as a means of probing the magnetosphere, and more recently as a means of actively controlling spacecraft potential. Experimentation in these areas has proven valuable, yet at times confusing, due to the interaction of the electron beam with the ambient plasma. The OSS-1/STS-3 Mission in March 1982 provided a unique opportunity to study beam-plasma interactions at an altitude of 240 km. On board for this mission was a Fast Pulse Electron Generator (FPEG). Measurements made by the Plasma Diagnostics Package (PDP) while extended on the Orbiter RMS show modifications of the ion and electron energy distributions during electron beam injection. Observations made by charged particle detectors are discussed and related to measurements of Orbiter potential. Several of the PDP instruments, the joint PDP/FPEG experiment, and observations made during electron beam injection are described.

Paterson, W.↗
NASA NTRS2009

Launch Order, Launch Separation, and Loiter in the Constellation 1 1/2-Launch Solution

The NASA Constellation Program (CxP) is developing a two-element Earth-to-Orbit launch system to enable human exploration of the Moon. The first element, Ares I, is a human-rated system that consists of a first stage based on the Space Shuttle Program's solid rocket booster (SRB) and an upper stage that consists of a four-crew Orion capsule, a service module, and a Launch Escape System. The second element, Ares V, is a Saturn V-plus category launch system that consists of the core stage with a cluster of six RS-68B engines and augmented with two 5.5-segment SRBs, a Saturn-derived J-2X engine powering an Earth Departure Stage (EDS), and the lunar-lander vehicle payload, Altair. Initial plans called for the Ares V to be launched first, followed the next day by the Ares I. After the EDS performs the final portion of ascent and subsequent orbit circularization, the Orion spacecraft then performs a rendezvous and docks with the EDS and its Altair payload. Following checkout, the integrated stack loiters in low Earth orbit (LEO) until the appropriate Trans-Lunar Injection (TLI) window opportunity opens, at which time the EDS propels the integrated Orion Altair to the Moon. Successful completion of this 1 1/2-launch solution carries risks related to both the orbital lifetime of the assets and the probability of achieving the launch of the second vehicle within the orbital lifetime of the first. These risks, which are significant in terms of overall system design choices and probability of mission success, dictated a thorough reevaluation of the launch strategy, including the order of vehicle launch and the planned time period between launches. The goal of the effort described in this paper was to select a launch strategy that would result in the greatest possible expected system performance, while accounting for launch risks and the cost of increased orbital lifetime. Discrete Event Simulation (DES) model of the launch strategies was created to determine the probability of a second launch not occurring in a timely fashion (i.e., before the assets waiting in LEO expire). An overview of the launch strategy evaluation process is presented, along with results of specific cases that were analyzed. A high-level comparison of options is then presented, along with the conclusion derived from the analysis.

Stromgren, Chel↗
NASA NTRS1989

Orbit and attitude determination results during launch support operations for SBS-5

Presented are orbit and attitude determination results from the launch of Satellite Business Systems (SBS)-5 satellite on September 8, 1988 by Arianespace. SBS-5 is a (HS-376) spin stabilized spacecraft. The launch vehicle injected the spacecraft into a low inclination transfer orbit. Apogee motor firing (AMF) attitude was achieved with trim maneuvers. An apogee kick motor placed the spacecraft into drift orbit. Postburn, reorientation and spindown maneuvers were performed during the next 25 hours. The spacecraft was on-station 19 days later. The orbit and attitude were determined by both an extended Kalman filter and a weighted least squares batch processor. Although the orbit inclination was low and the launch was near equinox, post-AMF analysis indicated an attitude declination error of 0.034 deg., resulting in a saving of 8.5 pounds of fuel. The AMF velocity error was 0.4 percent below nominal. The post-AMF drift rate was determined with the filter only 2.5 hours after motor firing. The filter was used to monitor and retarget the reorientation to orbit normal in real time.

Hartman, K. R.↗
NASA NTRS1984

Space plasma investigations on the First Spacelab Mission

A series of investigations was carried out during the First Spacelab Mission to study the space plasma environment from the Shuttle/Spacelab. The emphasis of the experiments was on the performance of particle injections from the Shuttle/Spacelab and studies of the ensuing effects on the orbiter, the near orbiter environment, and the earth's atmosphere. Results of these experiments, including electron beam injections, plasma injections, neutral injections, electron-plasma injections, electron-neutral injections, and ambient environmental measurements, are reviewed.

Roberts, W. T.↗
NASA NTRS1997

Lunar Prospector Mission Design

The National Aeronautics and Space Administration (NASA) has selected Lunar Prospector as one of the discovery missions to conduct solar system exploration science investigations. The mission is NASA's first lunar voyage to investigate key science objectives since Apollo and will be launched in September 1997. In keeping with discovery program requirements to reduce total mission cost and utilize new technology, Lunar Prospector's mission design and control will focus on the use of innovative and proven trajectory analysis programs. As part of this effort, the Ames Research Center and the Goddard Space Flight Center have become partners in the Lunar Prospector trajectory team to provide the trajectory analysis and orbit determination support. The mission design of Lunar Prospector is characterized by a direct minimum energy transfer trajectory to the moon with an injection into a circular 100 km altitude polar mapping orbit. The mission design will incorporate the optimization of the reference trajectory for a new launch vehicle. An extended mission includes a periselene altitude approaching 10 km. As part of the navigation, a new lunar potential model will be used that incorporates data from the recent Clementine mission. This paper provides understanding of the design of the Lunar Prospector mission and the optimization process of the nominal trajectory, orbit determination, launch windows, and lunar orbit targets, and discusses the feasibility of meeting the mission goals.

Folta, David↗
NASA NTRS1999

Submillimeter Wave Astronomy Satellite (SWAS) Launch and Early Orbit Support Experiences

The Submillimeter Wave Astronomy Satellite (SWAS) was successfully launched on December 6, 1998 at 00:58 UTC. The two year mission is the fourth in the series of Small Explorer (SMEX) missions. SWAS is dedicated to the study of star formation and interstellar chemistry. SWAS was injected into a 635 km by 650 km orbit with an inclination of nearly 70 deg by an Orbital Sciences Corporation Pegasus XL launch vehicle. The Flight Dynamics attitude and navigation teams supported all phases of the early mission. This support included orbit determination, attitude determination, real-time monitoring, and sensor calibration. This paper reports the main results and lessons learned concerning navigation, support software, star tracker performance, magnetometer and gyroscope calibrations, and anomaly resolution. This includes information on spacecraft tip-off rates, first-day navigation problems, target acquisition anomalies, star tracker anomalies, and significant sensor improvements due to calibration efforts.

Kirschner, S.↗
NASA NTRS1999

Submillimeter Wave Astronomy Satellite (SWAS) Launch and Early Orbit Support Experiences

The Submillimeter Wave Astronomy Satellite (SWAS) was successfully launched on December 6, 1998 at 00:58 UTC. The two year mission is the fourth in the series of Small Explorer (SMEX) missions. SWAS is dedicated to the study of star formation and interstellar chemistry. SWAS was injected into a 635 km by 650 km orbit with an inclination of nearly 70 deg by an Orbital Sciences Corporation Pegasus XL launch vehicle. The Flight Dynamics attitude and navigation teams supported all phases of the early mission. This support included orbit determination, attitude determination, real-time monitoring, and sensor calibration. This paper reports the main results and lessons learned concerning navigation, support software, star tracker performance, magnetometer and gyroscope calibrations, and anomaly resolution. This includes information on spacecraft tip-off rates, first-day navigation problems, target acquisition anomalies, star tracker anomalies, and significant sensor improvements due to calibration efforts.

Kirschner, S.↗
DOE OSTI2026

Numerical assessment of the impact of the guiding-centre approximation on fast ion simulations in NSTX

Guiding-centre (GC) and full-orbit (FO) simulations of the beam-injected fast ion distribution and the corresponding neutron emissivity have been carried out for magnetohydrodynamics-quiescent National Spherical Torus eXperiment (NSTX) plasmas, using a combination of ASCOT5 and DRESS, to assess the suitability of the GC approximation for fast ions in NSTX. It was found that GC and FO simulations predicted substantially different steady-state distributions in both position and velocity space and different neutron emissivity profiles, leading to a 15% reduction in the predicted global neutron rate for FO relative to GC. These changes accompany a higher magnetic moment in FO, and correspond to a change in particle orbits from co-passing to trapped and stagnation orbits. ASCOT5 was also benchmarked against TRANSP/NUBEAM with input loaded entirely from TRANSP/NUBEAM output files, with agreement found between the GC simulations when finite Larmor radius (FLR) corrections were omitted. ASCOT5 FO and TRANSP/NUBEAM with FLR produced fast ion distributions which differed in localised regions, but predicted global neutron rates which agree within 3%.

ASCOT↗
NASA NTRS1994

Project ECHO: Electronic Communications from Halo Orbit

The design of a communications relay to provide constant access between the Earth and the far side of the Moon is presented. Placement of the relay in a halo orbit about the L2 Earth-Moon Lagrange point allows the satellite to maintain constant simultaneous communication between Earth and scientific payloads on the far side of the Moon. The requirements of NASA's Discovery-class missions adopted and modified for this design are: total project cost should not exceed $150 million excluding launch costs, launch must be provided by Delta-class vehicle, and the satellite should maintain an operational lifetime of 10 to 15 years. The spacecraft will follow a transfer trajectory to the L2 point, after launch by a Delta II 7925 vehicle in 1999. Low-level thrust is used for injection into a stationkeeping-free halo orbit once the spacecraft reaches the L2 point. The shape of this halo orbit is highly elliptical with the maximum excursion from the L2 point being 35000 km. A spun section and despun section connected through a bearing and power transfer assembly (BAPTA) compose the structure of the spacecraft. Communications equipment is placed on the despun section to provide for a stationary dual parabolic offset-feed array antenna system. The dual system is necessary to provide communications coverage during portions of maximum excursion on the halo orbit. Transmissions to the NASA Deep Space Network 34 m antenna include six channels (color video, two voice, scientific data from lunar payloads, satellite housekeeping and telemetry and uplinked commands) using the S- and X-bands. Four radioisotope thermoelectric generators (RTG's) provide a total of 1360 W to power onboard systems and any two of the four Hughes 13 cm ion thrusters at once. Output of the ion thrusters is approximately 17.8 mN each with xenon as the propellant. Presence of torques generated by solar pressure on the antenna dish require the addition of a 'skirt' extending from the spun section of the satellite for balance. Total mass of the satellite is approximately 900 kg at a cost of $130 million FY99.

Borrelli, Jason↗
NASA NTRS1996

Analysis of Apollo 8 Photography and Visual Observations

Apollo 8 was launched from Cape Kennedy, Fla., at 7 :50 a.m., e.s.t., on December 21, 1968. Two hours 50 minutes later, translunar injection was performed; and astronauts Col. Frank Borman, the commander; Capt. James A. Lovell, Jr., the command module pilot; and Maj. (now Lt. Col.) William A. Anders, the lunar module pilot, were on their way to the Moon. The spacecraft was placed in an elliptical lunar orbit at 69 hours 8 minutes after liftoff. After flying two elliptical orbits of 168.5 by 60 nautical miles with an inclination of 12° to the Equator, the spacecraft was placed in a nearly circular orbit of 59.7 by 60.7 nautical miles, in which it remained for eight orbits. At 89 hours 19 minutes, trans earth injection was performed from behind the Moon. A nearly flawless mission was completed on the morning of December 27 when splashdown occurred in the Pacific Ocean after a total elapsed time of 147 hours. Lt. Gen. Sam C. Phillips, the Director of the Apollo Program, announced that such a mission was being considered at a press conference on August 19, 1968. Formal announcement that NASA was preparing Apollo 8 for an orbital flight around the Moon was released to the press on November 12, 1968. The primary purpose of this mission was to further progress toward the goal of landing men on the Moon by gaining operational experience and testing the Apolio systems. However, a great effort was also made to accomplish worthwhile scientific tasks with photography and visual observations by the astronauts. In planning the scientific tasks to be attempted on this mission, it was obvious that one of the prime tasks should be photography of the lunar surface. Such photography would furnish valuable information on the following : 1. Approach topography and landmarks for the early Apollo landings 2. The scientific merit and the roughness of areas for possible follow-on Apollo landings 3. The broad structure and characteristics of the lunar surface During the orbital part of the mission, a major portion of the lunar far side would be in sunlight. Although almost all of the far side of the Moon has been photographed by the automated Lunar Orbiter spacecraft, the photography generally was made with the spacecraft relatively far from the Moon, limiting the Lunar Orbiter photographs to an average resolution of approximately 100 meters. Thus, Apollo photographs of the far side would have much better resolution than existing pictures. Finally, it was recognized that contamination, both as it relates to window fogging (which did occur) and to contamination clouds around t he spacecraft, should be studied for both scientific and operational interests.

Richard J. Allenby↗
NASA NTRS1984

Commercial launch vehicles and upper stages

Since the beginning of the space age in October 1957, a family of expendable launch vehicles, capable of launching a wide range of payloads, was developed along with the Space Shuttle and a number of upper stages. A brief description is presented of selected orbits which have proved to be most useful for initial or conceptual understanding of space operations, taking into account direct injection and Hohman transfers, and synchronous and sun-synchronous orbits. Early American boosters are discussed along with current expendable launch vehicles, giving attention to the Vanguard, Redstone and Juno, Saturn 1B and Saturn V, Scout, the Atlas booster, Atlas Centaur, Delta, Titan IIIC, and Ariane. Details regarding the Space Shuttle are considered along with PAM-D, PAM-A, PAM-DII, TOS, IUS, Centaur-G, and Syncom-IV and Intelsat-VI.

Mahon, J.↗
NASA NTRS1972

Apollo 16 Press Kit

The Apollo 16 spacecraft is scheduled for launch on Apr. 16, 1972 from Complex 39A at the Kennedy Space Center, Florida by the Saturn V launch vehicle. Crewmen are mission commander John W. Young, command module pilot Thomas K. Mattingly II and lunar module pilot Charles M. Duke Jr. Objectives of the mission, to last up to 12 days, as outlined by NASA: to perform selenological inspection, survey and sampling of materials in a preselected region of Descartes using a lunar roving' vehicle; deploy and activate Apollo surface experiments; develop man's capability to work in the lunar environment; obtain photographs of candidate exploration sites; and toconduct inflight experiments and photographic tasks in lunar orbit. Following launch, the spacecraft will reach Earth Parking Orbit and remain in orbit for about two and one-half revolutions prior to Translunar Injection. Next, the Command and Service Module docks with the Lunar Module and the spacecraft "coasts" to the moon. In orbit around the moon, the Command and Service Module/Lunar Module combination will descend to within 50,000 feet of the lunar surface before undocking. The Lunar Module will continue to descend while the Command and Service Module returns to an orbit approximately 60 miles high. Stay time on the lunar surface is scheduled for approximately 73 hours. The ascent stage of the Lunar Module then lifts the astronauts back into lunar orbit where they will dock with the Command/Service Module. The Lunar Module is jettisoned and Transearth Injection follows. Just prior to reentry into the earth's atmosphere, the Service Module is jettisoned, and the astronauts in the Command Module splashdown in the Pacific Ocean. The target point for end-of-mission splashdown is at 05 degrees 0 minutes north latitude and 158 degrees 40 minutes west longitude or approximately 985 nautical miles south of Honolulu, Hawaii. Splashdown is scheduled for Apr. 28, 1972 at 10:30 a.m. Hawaiian Standard Time (2:30 p.m. CST). Recovery forces for Apollo 16, stationed in both the Atlantic and Pacific Oceans, will consist of three ships, nine aircraft and nearly 1,700 personnel. CTF-130 (Manned Spacecraft Recovery Force, Pacific) forces will be stationed south of Hawaii. Three ships, eight helicopters and three Air Force HC-130H aircraft, and nearly 1,100 personnel, will take part. Task Force 140 (Manned Spacecraft Recovery Force, Atlantic), comprising one ship, six HC-130H aircraft, three helicopters and approximately 300 personnel, will be positioned for possible launch abort operations. Two ships in the Atlantic will also be used for acoustical testing. Other forces, primarily aircraft and personnel of the Air Force Aerospace Rescue and Recovery Service will be on alert around the world for contingency recovery support.

Source record↗
NASA NTRS1995

Robotic Planetary Science Missions Enabled with Small NTR Engine/Stage Technologies

The high specific impulse (Isp) and engine thrust-to-weight ratio of liquid hydrogen (LH2)-cooled nuclear thermal rocket (NTR) engines makes them ideal for upper stage applications to difficult robotic planetary science missions. A small 15 thousand pound force (klbf) NTR engine using a uranium-zirconium-niobium 'ternary carbide' fuel (Isp approximately 960 seconds at approximately 3025K) developed in the Commonwealth of Independent States (CIS) is examined and its use on an expendable injection stage is shown to provide major increases in payload delivered to the outer planets (Saturn, Uranus, Neptune and Pluto). Using a single 'Titan IV-class' launch vehicle, with a lift capability to low Earth orbit (LEO) of approximately 20 metric tons (t), an expendable NTR upper stage can inject two Pluto 'Fast Flyby' spacecraft (PFF/SC) plus support equipment-combined mass of approximately 508 kg--on high energy, '6.5-9.2 year' direct trajectory missions to Pluto. A conventional chemical propulsion mission would use a liquid oxygen (LOX)/LH2 'Centaur' upper stage and two solid rocket 'kick motors' to inject a single PFF/SC on the same Titan IV launch vehicle. For follow on Pluto missions, the NTR injection stage would utilize a Jupiter 'gravity assist' (JGA) maneuver to launch a LOX/liquid methane (CH4) capture stage (Isp approximately 375 seconds) and a Pluto 'orbiter' spacecraft weighing between approximately 167-312 kg. With chemical propulsion, a Pluto orbiter mission is not a viable option because c inadequate delivered mass. Using a 'standardized' NTR injection stage and the same single Titan IV launch scenario, 'direct flight' (no gravity assist) orbiter missions to Saturn, Uranus and Neptune are also enabled with transit times of 2.3, 6.6, and 12.6 years, respectively. Injected mass includes a storable, nitrogen tetroxide/monomethyl hydrazine (N2O4/MMH) capture stage (Isp approximately 330 seconds) and orbiter payloads 340 to 820% larger than that achievable using a LOX/LH2-fueled injection stage. The paper discusses NTR technology and mission characteristics, shows NTR stage and payload accommodations within the 26.2 m long Titan IV payload fairing, and discusses NTR stage performance as a function of assumed cryogenic tank technology.

Borowski, Stanley K.↗
NASA NTRS1972

A common solar electric upper stage for multi-mission applications.

The concept of using a solar-electric upper stage for multi-mission application is practical, economical, and versatile approach to space expMission capabilities include a Mercury orbiter, asteroid and comet rendezvous, a solar probe, and extra-ecliptic probe, and outer planet flybys and orbiters. The 17.5-kw vehicle has an injected mass of 1500 to 2500 kg and carries up to 500 kg of attached or separable payload packages. Mission characteristics, performance tradeoffs and a concepegn configuration are presented, and program plans, cost estimates and advanced technology requirements are identified.

Meissinger, H. F.↗
NASA NTRS1972

An investigation to improve selenodetic control on the lunar limb utilizing Apollo 15 trans-earth photography

A study is made using actual metric photography taken by Apollo 15 after the trans earth injection when the spacecraft left the lunar orbit. After measurements on 12 frames were reduced for 11 known control and unknown points on the eastern limb, a least squares adjustment program provided the simultaneous solution for the selected points. The results indicate an improvement in selenodetic control may be achieved over a limited portion of the lunar limb; however, the solution could be further improved by strengthening the geometry of intersecting rays through additional observations of the same area using film from the succeeding Apollo missions.

Riotte, W.↗
NASA NTRS1986

Mars orbit selection

Parking orbits for a manned Mars mission are examined for ease of access to the Martian moons. Delta V plots for a variety of burns versus elliptical orbit apoapsis are included. A high elliptical orbit (24 hour period, 500 km periapsis, 20 to 30 deg. inclination) minimizes delta V to the Martian moons and Mars orbit insertion (MOI) and trans-Earth injection (TEI) delta Vs.

Gus R Babb↗