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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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This video gives an overview of work being done by the Space Propulsion Technology Division at LeRC. This division conducts research on chemical, nuclear-thermal, and solar propulsion systems and propellants. Two ongoing projects highlighted are a low-thrust rocket for moving around in Earth orbit and large unmanned cargo rockets, both for use with the Space Station.
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.
This paper details a numerical investigation of performance losses in low-thrust solar thermal rocket nozzles. The effects of nozzle geometry on three types of losses were studied; finite rate dissociation-recombination kinetic losses, two dimensional axisymmetric divergence losses, and compressible viscous boundary layer losses. Short nozzle lengths and supersonic flow produce short residence times in the nozzle and a nearly frozen flow, resulting in large kinetic losses. Variations in geometry have a minimal effect on kinetic losses. Divergence losses are relatively small, and careful shaping of the nozzle can nearly eliminate them. The boundary layer in these small nozzles can grow to a major fraction of nozzle radius, and cause large losses. These losses are attributed to viscous drag on the nozzle walls and flow blockage by the boundary layer, especially in the throat region. Careful shaping of the nozzle can produce a significant reduction in viscous losses.
This volume contains papers that have been accepted for presentation at the Workshop. Topics considered include: On the sample return from Vesta by low-thrust spacecraft; Astronomical evidence linking Vesta to the HED meteorites; Geologic mapping of Vesta with the Hubble Space Telescope; A space mission to Vesta; Asteroid spectroscopy; The thermal history of asteroid 4 Vesta, based on radionuclide and collision heating; Mineralogical records of early planetary processes on Vesta.
Radiation-cooled rockets are used for a range of low-thrust propulsion functions, including apogee insertion, attitude control, and repositioning of satellites, reaction control of launch vehicles, and primary propulsion for planetary space- craft. The key to high performance and long lifetimes for radiation-cooled rockets is the chamber temperature capability. The material system that is currently used for radiation-cooled rockets, a niobium alloy (C103) with a fused silica coating, has a maximum operating temperature of 1370 C. Temperature limitations of C103 rockets force the use of fuel film cooling, which degrades rocket performance and, in some cases, imposes a plume contamination issue from unburned fuel. A material system composed of a rhenium (Re) substrate and an iridium (Ir) coating has demonstrated operation at high temperatures (2200 C) and for long lifetimes (hours). The added thermal margin afforded by iridium-coated rhenium (Ir/Re) allows reduction or elimination of fuel film cooling. This, in turn, leads to higher performance and cleaner spacecraft environments. There are ongoing government- and industry-sponsored efforts to develop flight Ir/ Re engines, with the primary focus on 440-N, apogee insertion engines. Complementing these Ir/Re engine development efforts is a program to address specific concerns and fundamental characterization of the Ir/Re material system, including (1) development of Ir/Re rocket fabrication methods, (2) establishment of critical Re mechanical properly data, (3) development of reliable joining methods, and (4) characterization of Ir/Re life-limiting mechanisms.
This paper presents several mission designs for heliospheric boundary exploration using spacecraft with low-thrust ion engines as the primary mode of propulsion The mission design goal is to transfer a 200-kg spacecraft to the heliospheric boundary in minimum time. The mission design is a combined trajectory and propulsion system optimization problem. Trajectory design variables include launch date, launch energy, burn and coast arc switch times, thrust steering direction, and planetary flyby conditions. Propulsion system design parameters include input power and specific impulse. Both SEP and NEP spacecraft arc considered and a wide range of launch vehicle options are investigated. Numerical results are presented and comparisons with the all chemical heliospheric missions from Ref 9 are made.
An analysis was made to determine the maximum powerplant thrust-to-weight ratio possible with a single-cavity vortex gaseous reactor in which all the hydrogen propellant must diffuse through a fuel-rich region. An assumed radial temperature profile was used to represent conduction, convection, and radiation heat-transfer effects. The effect of hydrogen property changes due to dissociation and ionization was taken into account in a hydrodynamic computer program. It is shown that, even for extremely optimistic assumptions of reactor criticality and operating conditions, such a system is limited to reactor thrust-to-weight ratios of about 1.2 x 10(exp -3) for laminar flow. For turbulent flow, the maximum thrust-to-weight ratio is less than 10(exp -3). These low thrusts result from the fact that the hydrogen flow rate is limited by the diffusion process. The performance of a gas-core system with a specific impulse of 3000 seconds and a powerplant thrust-to-weight ratio of 10(exp -2) is shown to be equivalent to that of a 1000-second advanced solid-core system. It is therefore concluded that a single-cavity vortex gaseous reactor in which all the hydrogen must diffuse through the nuclear fuel is a low-thrust device and offers no improvement over a solid-core nuclear-rocket engine. To achieve higher thrust, additional hydrogen flow must be introduced in such a manner that it will by-pass the nuclear fuel. Obviously, such flow must be heated by thermal radiation. An illustrative model of a single-cavity vortex system employing supplementary flow of hydrogen through the core region is briefly examined. Such a system appears capable of thrust-to-weight ratios of approximately 1 to 10. For a high-impulse engine, this capability would be a considerable improvement over solid-core performance. Limits imposed by thermal radiation heat transfer to cavity walls are acknowledged but not evaluated. Alternate vortex concepts that employ many parallel vortices to achieve higher hydrogen flow rates offer the possibility of sufficiently high thrust-to-weight ratios, if they are not limited by short thermal-radiation path lengths.
This paper discusses the (chemical or electric) propulsion system requirements necessary to increase the Spartan Lite science mission lifetime to over a year. Spartan Lite is an extremely low-cost (less than 10 M) spacecraft bus being developed at the NASA Goddard Space Flight Center to accommodate sounding rocket class (40 W, 45 kg, 35 cm dia by 1 m length) payloads. While Spartan Lite is compatible with expendable launch vehicles, most missions are expected to be tertiary payloads deployed by. the Space Shuttle. To achieve a one year or longer mission life from typical Shuttle orbits, some form of propulsion system is required. Chemical propulsion systems (characterized by high thrust impulsive maneuvers) and electrical propulsion systems (characterized by low-thrust long duration maneuvers and the additional requirement for electrical power) are discussed. The performance of the Spartan Lite attitude control system in the presence of large disturbance torques is evaluated using the Trectops(Tm) dynamic simulator. This paper discusses the performance goals and resource constraints for candidate Spartan Lite propulsion systems and uses them to specify quantitative requirements against which the systems are evaluated.
The Radiation and Technology Demonstration (RTD) Mission has the primary objective of demonstrating high-power (10 kilowatts) electric thruster technologies in Earth orbit. This paper discusses the conceptual design of the RTD spacecraft photovoltaic (PV) power system and mission performance analyses. These power system studies assessed multiple options for PV arrays, battery technologies and bus voltage levels. To quantify performance attributes of these power system options, a dedicated Fortran code was developed to predict power system performance and estimate system mass. The low-thrust mission trajectory was analyzed and important Earth orbital environments were modeled. Baseline power system design options are recommended on the basis of performance, mass and risk/complexity. Important findings from parametric studies are discussed and the resulting impacts to the spacecraft design and cost.
A closed-loop system for the demonstration of autonomous satellite formation flying technologies using hardware-in-the-loop has been developed. Making use of a GPS signal simulator with a dual radio frequency outlet, the system includes two GPS space receivers as well as a powerful onboard navigation processor dedicated to the GPS-based guidance, navigation, and control of a satellite formation in real-time. The closed-loop system allows realistic simulations of autonomous formation flying scenarios, enabling research in the fields of tracking and orbit control strategies for a wide range of applications. The autonomous closed-loop formation acquisition and keeping strategy is based on Lyapunov's direct control method as applied to the standard set of Keplerian elements. This approach not only assures global and asymptotic stability of the control but also maintains valuable physical insight into the applied control vectors. Furthermore, the approach can account for system uncertainties and effectively avoids a computationally expensive solution of the two point boundary problem, which renders the concept particularly attractive for implementation in onboard processors. A guidance law has been developed which strictly separates the relative from the absolute motion, thus avoiding the numerical integration of a target trajectory in the onboard processor. Moreover, upon using precise kinematic relative GPS solutions, a dynamical modeling or filtering is avoided which provides for an efficient implementation of the process on an onboard processor. A sample formation flying scenario has been created aiming at the autonomous transition of a Low Earth Orbit satellite formation from an initial along-track separation of 800 m to a target distance of 100 m. Assuming a low-thrust actuator which may be accommodated on a small satellite, a typical control accuracy of less than 5 m has been achieved which proves the applicability of autonomous formation flying techniques to formations of satellites as close as 50 m.
Conceptual in-space transfer stages, including those utilizing solar electric propulsion, chemical propulsion, and chemical propulsion with aerobraking or aerocapture assist at Mars, were evaluated. Roundtrip Mars sample return mission vehicles were analyzed to determine how specific system technology selections influence payload delivery capability. Results show how specific engine, thruster, propellant, capture mode, trip time and launch vehicle technology choices would contribute to increasing payload or decreasing the size of the required launch vehicles. Heliocentric low-thrust trajectory analyses for Solar Electric Transfer were generated with the SEPTOP code.
NASA Lewis Research Center's On-Board Propulsion program (OBP) is developing low-thrust chemical propulsion technologies for both satellite and vehicle reaction control applications. There is a vigorous international competition to develop new, highperformance bipropellant engines. High-leverage bipropellant systems are critical to both commercial competitiveness in the international communications market and to cost-effective mission design in government sectors. To significantly improve bipropellant engine performance, we must increase the thermal margin of the chamber materials. Iridium-coated rhenium (Ir/Re) engines, developed and demonstrated under OBP programs, can operate at temperatures well above the constraints of state-of-practice systems, providing a sufficient margin to maximize performance with the hypergolic propellants used in most satellite propulsion systems.
There is increasing interest in employing Solar Electric Propulsion (SEP) for new missions requiring transfer from low Earth orbit to the Earth-Moon Lagrange point, L1. Mission architecture plans place the Gateway Habitat at L1 in the 2011 to 2016 timeframe. The Gateway Habitat is envisioned to be used for Lunar exploration, space telescopes, and planetary mission staging. In these scenarios, an SEP stage, or "tug," is used to transport payloads to L1--such as the habitat module, lunar excursion and return vehicles, and chemical propellant for return crew trips. SEP tugs are attractive because they are able to efficiently transport large (less than 10,000 kg) payloads while minimizing propellant requirements. To meet the needs of these missions, a preliminary conceptual design for a general-purpose SEP tug was developed that incorporates several of the advanced space power and in-space propulsion technologies (such as high-power gridded ion and Hall thrusters, high-performance thin-film photovoltaics, lithium-ion batteries, and advanced high-voltage power processing) being developed at the NASA Glenn Research Center. A spreadsheet-based vehicle system model was developed for component sizing and is currently being used for mission planning. This model incorporates a low-thrust orbit transfer algorithm to make preliminary determinations of transfer times and propellant requirements. Results from this combined tug mass estimation and orbit transfer model will be used in a higher fidelity trajectory model to refine the analysis.
Generic algorithms have gained popularity as an effective procedure for obtaining solutions to traditionally difficult space mission optimization problems. In this paper, a brief survey of the use of genetic algorithms to solve astrodynamics problems is presented and is followed by new results obtained from applying a Pareto genetic algorithm to the optimization of low-thrust interplanetary spacecraft missions.
NASA's proposed mission to send a single electric propulsion spacecraft to orbit Callisto, Ganymeds, and Europa within a decade would require a very complex trajectory. Strong multi-body effects combined with low-thrust control of capture and escape will make the trajectory challenging. This paper describes an optimal trajectory that begins in low Earth orbit and ends in low Io orbit.
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With the complexities in computing optimal low thrust trajectories, easily-computed, good sub-optimal trajectories provide both a practical alternative for mission designers and a starting point for optimisation. The present paper collects in one place for easy reference and comparison several exact solutions that have been obtained in the literature over the last few decades: the logarithmic spiral, Pinkham's variant thereof, Forbes spiral, the exponential sinusoid, the case of constant radial thrust, Markopoulos's Keplerian thrust arcs, Lawden's spiral, and the analogous Bishop and Azimov spiral.
In situ analysis of a cometary nucleus and return of a sample are high priority scientific goals.