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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 379 records · Page 21
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NASA NTRS2015

NASA's Space Launch System Program Update

Hardware and software for the world's most powerful launch vehicle for exploration is being welded, assembled, and tested today in high bays, clean rooms and test stands across the United States. NASA's Space Launch System (SLS) continued to make significant progress in the past year, including firing tests of both main propulsion elements, manufacturing of flight hardware, and the program Critical Design Review (CDR). Developed with the goals of safety, affordability, and sustainability, SLS will deliver unmatched capability for human and robotic exploration. The initial Block 1 configuration will deliver more than 70 metric tons (t) (154,000 pounds) of payload to low Earth orbit (LEO). The evolved Block 2 design will deliver some 130 t (286,000 pounds) to LEO. Both designs offer enormous opportunity and flexibility for larger payloads, simplifying payload design as well as ground and on-orbit operations, shortening interplanetary transit times, and decreasing overall mission risk. Over the past year, every vehicle element has manufactured or tested hardware, including flight hardware for Exploration Mission 1 (EM-1). This paper will provide an overview of the progress made over the past year and provide a glimpse of upcoming milestones on the way to a 2018 launch readiness date.

May, Todd↗
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 NTRS1966

Prospects for manned Mars missions

Studies called "EMPIRE" (standing for Early Manned Planetary-Interplanetary Roundtrip Expeditions) were initiated early in 1962 by the Future Projects Office for the Marshall Space Flight Center. These and related studies by other NASA centers were intended to bring to light the technical problems associated with such missions. The emphasis of these studies was on single and multiple planetary fty-bys, with some consideration given to planetary orbital and landing missions. The time period considered was the early 1970's. The contractors were Aeronutronic Division (Philco-Ford Motor Company), Lockheed Missiles and Space Company and General Dynamics/Astronautics. Studies relating to the time period 1975-85 were initiated in early 1963, by contracts with General Dynamics/ Fort Worth and Douglas Aircraft Company. Follow-on EMPIRE studies were carried out by Lockheed and General Dynamics/Astronautics. In addition, studies were made of a Mars Excursion Module by Aeronutronic, a Mars Mission Module (MMM) by North American Aviation's Space and Information Division, and an Earth Return Module (ERM) by Lockheed under the direction of the Manned Spacecraft Center. Complementary to these efforts were studies initiated by the Ames Research Center early in 1963 for a Manned Mars Landing and Return Mission. These studies were conducted by North American Aviation and by the TRW Space Technology Laboratories.

MARS /PLANET/↗
NASA NTRS1980

Safety of liquid propulsion systems for STS payloads

STS payload safety requirements for Earth-storable liquid propulsion systems are reviewed with respect to applicable safety documents. The effect of normal flight and emergency landing/abort launch profiles on propulsion system design is delineated. Spacecraft studied include Earth satellites, interplanetary spacecraft, and orbit transfer vehicles. Safety considerations for monopropellant (hydrazine), bipropellant (nitrogen tetroxide/monomethyl hydrazine), and ion (mercury) propulsion systems are described. Safety requirements for specific components of propulsion systems are reported. Analysis techniques for STS payload fracture control requirements for pressure vessels and critical support structure are discussed.

Lewis, J. C.↗
NASA NTRS1982

Space transportation

The physical laws governing launch vehicles are reviewed, together with the performance of launch vehicles thus far used and expected developments in space transportation systems. Launches are generally made in the direction of earth rotation to take advantage of the initial velocity in that direction, which can amount to about 0.46 km/sec at the equator. Spacecraft usually attain a parking orbit, where spacecraft conditions are assessed before moving on to final orbit or interplanetary trajectory. Rocket batteries were first used in a battle near Beijing in 1232, and manned space flight began with the Vostok flight in 1961, followed by the Apollo lunar landings in the later 1960s. Rocket thrust performance is analyzed, together with the thrust/mass ratio for ascent and descent. The Shuttle, capable of placing 29.5 t in LEO, will be or is equipped to also carry IUS and PAM engines for transferring payloads from LEO to GEO. Parallel burn boosters may be added to increase the payload capability, and Shuttle-derived launch vehicles may be developed to carry construction materials to space and return to earth for a runway landing. Alternatively, the Shuttle engines may be modularized in order to develop a heavy-lift launch vehicle for unmanned cargo ascents.

Davis, H.↗
NASA NTRS1984

Solar modulation of cosmic ray electrons 1978-1983

The ISEE 3 spacecraft has provided over 4 years of continuous data from nearby interplanetary space. By using these data together with prelaunch calibration data and precise electron spectra from balloon flights in 1979 and 1982 at Thompson, Manitoba, Canada, the flux level of approximately 1 GeV cosmic ray electrons with 1 month time resolution is monitored. The step-like onset of modulation observed for positive particles in 1979 is also seen in electrons, confirming that the dominant modulation mechanism is not sensitive to the charge of the particle. Hysteresis effects observed in the previous solar maximum by Burger and Swanenburg (1973) are not reproduced at the present solar maximum. A return to the relatively higher electron/proton ratios of the 1960s is a strong possibility. Conclusions must be drawn with caution, however, since the overall time structure of the present solar maximum is significantly different from that of the previous one.

Evenson, P.↗
NASA NTRS1988

Data Transport Subsystem - The SFOC glue

The design and operation of the Data Transport Subsystem (DTS) for the JPL Space Flight Operation Center (SFOC) are described. The SFOC is the ground data system under development to serve interplanetary space probes; in addition to the DTS, it comprises a ground interface facility, a telemetry-input subsystem, data monitor and display facilities, and a digital TV system. DTS links the other subsystems via an ISO OSI presentation layer and an LAN. Here, particular attention is given to the DTS services and service modes (virtual circuit, datagram, and broadcast), the DTS software architecture, the logical-name server, the role of the integrated AI library, and SFOC as a distributed system.

Parr, Stephen J.↗
NASA NTRS2003

Measurement of Characteristics of Micron Size Individual Dust Particles of Astrophysical Interest

A laboratory facility for levitating single isolated dust particles in an electrodynamic balance has been developed at NASA's Marshall Space Flight Center for conducting studies of the physical and optical properties of the analogs of interstellar and interplanetary dust grains of 0.2-20 micron size under controlled pressures/temperatures simulating astrophysical environments. We plan three classes of experiments using this facility: (1) Charge characteristics of micron size single dust grains: The photoelectric efficiencies, yields, and equilibrium potentials when exposed to UV radiation found from these measurements will provide much-needed photoelectric emission data for individual dust particles; (2) Infrared optical properties of dust particles: Specifically, we will determines the complex refractive indices, the extinction coefficients, the scattering phase functions, and the polarization properties of single dust grains of interest in interstellar environments, in the 1- 25 micron spectral region; (3) Condensation experiments to investigate the deposition of volatile gases on colder nucleated particles in dense interstellar clouds and lower planetary atmospheres: The measured data will permit determination of the sticking efficiencies of volatile gases of astrophysical interest. Brief descriptions of the experimental setup for the last two classes of experiments will be given. We will present results of measurements of photoelectric emission using 0.2-6.6 micron size silica particles exposed to UV radiation at 120-200 nm and also results of radiation pressure measurements using the same size silica particles and laser light at 5320 Angstrom.

Craven, P. D.↗
NASA NTRS2013

Flight Performance of the Inflatable Reentry Vehicle Experiment 3

The Inflatable Reentry Vehicle Experiment 3 (IRVE-3) launched July 23, 2012, from NASA Wallops Flight Facility (WFF) on a Black Brant XI suborbital sounding rocket and successfully performed its mission, demonstrating the survivability of a hypersonic inflatable aerodynamic decelerator (HIAD) in the reentry heating environment and also illustrating the effect of an offset center of gravity on the HIAD's lift-to-drag ratio. IRVE-3 was a follow-on to 2009's IRVE-II mission, which demonstrated exo-atmospheric inflation, reentry survivability - without significant heating - and the aerodynamic stability of a HIAD down to subsonic flight conditions. NASA Langley Research Center is leading the development of HIAD technology for use on future interplanetary and Earth reentry missions.

Dillman, Robert↗
NASA NTRS1990

Xenon ion propulsion for orbit transfer

For more than 30 years, NASA has conducted an ion propulsion program which has resulted in several experimental space flight demonstrations and the development of many supporting technologies. Technologies appropriate for geosynchronous stationkeeping, earth-orbit transfer missions, and interplanetary missions are defined and evaluated. The status of critical ion propulsion system elements is reviewed. Electron bombardment ion thrusters for primary propulsion have evolved to operate on xenon in the 5 to 10 kW power range. Thruster efficiencies of 0.7 and specific impulse values of 4000 s were documented. The baseline thruster currently under development by NASA LeRC includes ring-cusp magnetic field plasma containment and dished two-grid ion optics. Based on past experience and demonstrated simplifications, power processors for these thrusters should have approximately 500 parts, a mass of 40 kg, and an efficiency near 0.94. Thrust vector control, via individual thruster gimbals, is a mature technology. High pressure, gaseous xenon propellant storage and control schemes, using flight qualified hardware, result in propellant tankage fractions between 0.1 and 0.2. In-space and ground integration testing has demonstrated that ion propulsion systems can be successfully integrated with their host spacecraft. Ion propulsion system technologies are mature and can significantly enhance and/or enable a variety of missions in the nation's space propulsion program.

Rawlin, V. K.↗
NASA NTRS1973

Variations of the occurrence rate of discontinuities in the interplanetary magnetic field

A statistical study of magnetic field discontinuities for the interval from February to October 1968 using data from the NASA Goddard Space Flight Center/University of Rome experiment on Pioneer 8 has been conducted. High-resolution data (10-sec averages) have been employed, stringent criteria being utilized to select more than 16,000 events, about 50 per cent being identified as disturbances that are probably tangential discontinuities (TD). Near 1 AU the average occurrence rate of discontinuities is about 3.6/hr, whereas that of tangential-like discontinuities is about 1.6/hr. The rate of occurrence of discontinuities is found to be correlated with the directional change of the field across the discontinuity surface. The normal to the TD discontinuity plane is most often perpendicular to the Archimedean spiral field configuration. A decrease in the rate of occurrence of TD is found with increasing heliocentric distance and/or with decreasing heliocentric latitude. Possible interpretations of the effect are discussed.

Mariani, F.↗
NASA NTRS1999

A Flight Demonstration of Plasma Rocket Propulsion

The Advanced Space Propulsion Laboratory at the Johnson Space Center has been engaged in the development of a magneto-plasma rocket for several years. This type of rocket could be used in the future to propel interplanetary spacecraft. One advantageous feature of this rocket concept is the ability to vary its specific impulse so that it can be operated in a mode which maximizes propellant efficiency or a mode which maximizes thrust. This presentation will describe a proposed flight experiment in which a simple version of the rocket will be tested in space. In addition to the plasma rocket, the flight experiment will also demonstrate the use of a superconducting electromagnet, extensive use of heat pipes, and possibly the transfer of cryogenic propellant in space.

Petro, Andrew↗
NASA NTRSDate not supplied

Aerocapture Trajectory Design for Uranus Orbiter

Introduction: The recently released National Academies Planetary Science and Astrobiology Decadal Survey 2023-2032 [1] identified the Ice Giants as the top priority science destination. While the survey acknowledged the potential for either a Uranus Orbiter and Probe (UOP) mission or a Neptune-Triton Odyssey mission, it ultimately identified the former as the highest priority new flagship mission. UOP missions calls for a launch window of opportunity between 2031-2038 with 12-15 year interplanetary cruise time along with a fully-propulsive Uranus Orbit Insertion burn on the order of a few km/s. However, a mission to Uranus with the same science payload could utilize aerocapture for orbit insertion to achieve both a significant reduction in the interplanetary cruise time and reduction in propulsive burn costs. Why Aerocapture: Aerocapture is a promising propellant and time-saving orbital insertion technique for planetary destinations with an atmosphere. Although not flight-proven, previous aerocapture systems studies in the literature have demonstrated both the validity and robustness of the technique at various planetary destinations. With respect to the Ice Giant planets, Neptune has seen more of the analysis in the literature. For science missions at Neptune, aerocapture can enable 1.4 times more delivered mass to orbit than an all-propulsive mission for the same launch vehicle while reducing interplanetary cruise times by more than 3 years [2] Additionally with modern guidance and control, Neptune aerocapture with blunt-body aeroshells is realizable [3][4]. There are limited papers in the literature investigating Uranus aerocapture with those available providing a preliminary feasibility assessment [5]. Consequently, the two-year funded NASA Space Technology Mission Directorate (STMD)-funded project, titled Aerocapture System as an Enabling Technology for Ice Giants Missions, aims to mature the analysis and technology state of Uranus aerocapture. Trajectory Design: This paper presents the current state of the trajectory design in support of the new aerocapture project. The project design philosophy is inspired from recent Neptune aerocapture studies, which employed modern guidance and control, in the sense that blunt-body aeroshells are analyzed. An assessment of the theoretical flight path angle corridor width is conducted for a range of ballistic coefficients and lift-to-drag ratios for both Space Launch System and Falcon Heavy Launch Vehicle interplanetary trajectory solutions. The results from the corridor width assessment provide an assessment of the aerocapture design-space and qualitative metrics on trajectory design considerations. The Program to Optimize Simulated Trajectories II (POST2) is utilized to run Monte Carlo simulations of Uranus aerocapture three-degree-of-freedom bank angle modulated trajectories using a closed-loop numerical-predictor corrector guidance algorithm. UranusGRAM 2021 is utilized as the atmospheric model [6]. A Uranus-developed aerodatabase, originally derived from Mars Science Laboratory (MSL), is utilized to provide vehicle aerodynamics over a wide range of hypersonic flow regimes. A MSL-derived 70 deg 4.5m diameter sphere-cone aeroshell that houses the UOP payload mass is assumed. Robustness testing and performance analysis is conducted, including the assessment of entry state errors, atmosphere density variations, and aerodynamic dispersions. Post-aerocapture Delta-V and aerothermal statistics are formulated into propellant mass and TPS requirements. The results presented in the paper will demonstrate the trajectory viability of Uranus aerocapture. Preliminary Results: Preliminary trajectory design results indicates successful Uranus aerocapture with a blunt-body aeroshell housing the same payload mass as the UOP mission from an Earth-to-Uranus interplanetary trajectory arriving in less than 7 years. From this interplanetary trajectory, aerocapture provides an orbit insertion Delta-V capability of 6.9 km/s requiring less than 300 m/s for post-aerocapture correction burns (15% of wet mass allocated to propellant). To put this into perspective, the UOP study utilized an Earth-to-Uranus trajectory that arrives in 13 years and requires more than 1000 m/s for fully-propulsive orbit insertion (40% of wet mass allocated to propellant). Achieving the same 6.9 km/s Delta-V capability fully-propulsively is mass prohibitive (97% of wet mass allocated to propellant). Nevertheless, aerocapture has the potential to reduce interplanetary transit times to Uranus by half while delivering the same payload mass to orbit in a reduced propellant mass footprint.

Rohan Deshmukh↗
DOE OSTI2024

Star Truck : Interplanetary Bussing system

CubeSats are a standardized size of satellite. Used by elementary schools, universities, and hobbyists alike. CubeSats have made space exploration and research accessible to the general population. Utilizing CubeSats it is possible to make deep space exploration accessible as well. Using a flight path that takes advantage of gravitational assists and flybys we can use many forms of propulsion to get to Jupiter. However, to get farther nuclear power and propulsion can be utilized to bus hundreds of CubeSats at a time to interplanetary space. This craft is known as a Star Truck. The Star Truck will make interplanetary space assessable to the common man.

Belian, Olivia↗
NASA NTRS1973

A variation of the Davis-Smith method for in-flight determination of spacecraft magnetic fields.

A variation of a procedure developed by Davis and Smith (1968) is presented for the in-flight determination of spacecraft magnetic fields. Both methods take statistical advantage of the observation that fluctuations in the interplanetary magnetic field over short periods of time are primarily changes in direction rather than in magnitude. During typical solar wind conditions between 0.8 and 1.0 AU, a statistical analysis of 2-3 days of continuous interplanetary field measurements yields an estimate of a constant spacecraft field with an uncertainty of plus or minus 0.25 gamma in the direction radial to the sun and plus or minus 15 gammas in the directions transverse to the radial. The method is also of use in estimating variable spacecraft fields with gradients of the order of 0.1 gamma/day and less and in other special circumstances.

Belcher, J. W.↗
NASA NTRS2002

Trajectories for High Specific Impulse High Specific Power Deep Space Exploration

Flight times and deliverable masses for electric and fusion propulsion systems are difficult to approximate. Numerical integration is required for these continuous thrust systems. Many scientists are not equipped with the tools and expertise to conduct interplanetary and interstellar trajectory analysis for their concepts. Several charts plotting the results of well-known trajectory simulation codes were developed and are contained in this paper. These charts illustrate the dependence of time of flight and payload ratio on jet power, initial mass, specific impulse and specific power. These charts are intended to be a tool by which people in the propulsion community can explore the possibilities of their propulsion system concepts. Trajectories were simulated using the tools VARITOP and IPOST. VARITOP is a well known trajectory optimization code that involves numerical integration based on calculus of variations. IPOST has several methods of trajectory simulation; the one used in this paper is Cowell's method for full integration of the equations of motion. An analytical method derived in the companion paper was also evaluated. The accuracy of this method is discussed in the paper.

Polsgrove, Tara↗
NASA NTRS2005

Helium-3 Mining Aerostats in the Atmospheres of the Outer Planets

Imagine an interplanetary future where: a) d-He3 fusion produces most of Earth s energy needs without radioactivity or carbon emissions; b) Space transportation has been revolutionized by an efficient fusion propulsion system with exhaust velocity up to 0.088 c; c) Space commerce is stimulated by the existence of an interplanetary cargo worth $3-M a kilogram; and d) Unmanned probes travel to the nearest star systems with flight times less than a human lifetime.

VanCleve, Jeffrey E.↗