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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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Mathematical techniques used in minimum variance orbital navigation and guidance, midcourse phase, lunar parking, descent orbits, landing, ascent, and rendezvous
Go, no-go for Apollo based on orbital life time
Vapor replacement in propellant feed duct with subcooled liquid to meet engine starting requirement of space vehicle parked in orbit
Simplified guidance technique designed to put rocket in close orbit around Moon
Midcourse guidance, lunar parking and descent orbits, lunar landing phase, lunar ascent, and lunar and earth rendezvous
A high degree and order Lunar gravitational field is superimposed on the Earth-Moon Restricted Three Body model to capture the dominating forces on a spacecraft in the vicinity of the Moon. For the synchronously rotating Moon, periodic orbits in this model map repeat ground tracks and represent higher order solutions to the frozen orbit problem. The near-circular, stable or near-stable solutions are found over a wide range of defining characteristics making them suitable for long-lifetime parking applications such as science orbits, crew exploration vehicle parking orbits, and global coverage constellation orbits. A full ephemeris is considered for selected orbits to evaluate the validity of the time-invariant, simplified model. Of the most promising results are the low-altitude families of near-circular, inclined orbits that maintain long-term stability despite the highly non-spherical Lunar gravity. The method is systematic and enables rapid design and analysis of long-life orbits around any tidally-locked celestial body with an arbitrarily high degree and order spherical harmonic gravity field. .
The long-term stable trajectories around Europa, one of the Galilean moons of Jupiter, are analyzed for their potential applications in spacecraft trajectory design, such as end of mission desposal options, backup orbits, or intermediary targets for transfer trajectories.
Logistics system for synchronous orbit space stations supply, discussing parking to synchronous orbit transfer of reusable and disposable logistics vehicles
An actual geosynchronous Earth orbit-to-low Earth orbit (GEO-to-LEO) transfer is considered with reference to the aeroassisted flight experiment (AFE) spacecraft, and optimal trajectories are determined by minimizing the total characteristic velocity. The optimization is performed with respect to the time history of the controls (angle of attack and angle of bank), the entry path inclination and the flight time being free. Two transfer maneuvers are considered: direct ascent (DA) to LEO and indirect ascent (IA) to LEO via parking Earth orbit (PEO). By taking into account certain assumptions, the complete system can be decoupled into two subsystems: one describing the longitudinal motion and one describing the lateral motion. The angle of attack history, the entry path inclination, and the flight time are determined via the longitudinal motion subsystem. In this subsystem, the difference between the instantaneous bank angle and a constant bank angle is minimized in the least square sense subject to the specified orbital inclination requirement. Both the angles of attack and the angle of bank are shown to be constant. This result has considerable importance in the design of nominal trajectories to be used in the guidance of AFE and aeroassisted orbital transfer (AOT) vehicles.
An actual geosynchronous earth orbit-to-low earth orbit (GEO-to-LEO) transfer is considered with reference to the aeroassisted flight experiment (AFE) spacecraft, and optimal trajectories are determined by minimizing the total characteristic velocity. The optimization is performed with respect to the time history of the controls (angle of attack and angle of bank), the entry path inclination and the flight time being free. Two transfer maneuvers are considered: direct ascent (DA) to LEO and indirect ascent (IA) to LEO via parking earth orbit (PEO). By taking into account certain assumptions, the complete system can be decoupled into two subsystems: one describing the longitudinal motion and one describing the lateral motion. The angle of attack history, the entry path inclination, and the flight time are determined via the longitudinal motion subsystem. In this subsystem, the difference between the instantaneous bank angle and a constant bank angle is minimized in the least square sense subject to the specified orbital inclination requirement. Both the angles of attack and the angle of bank are shown to be constant. This result has considerable importance in the design of nominal trajectories to be used in the guidance of AFE and aeroassisted orbital transfer (AOT) vehicles.
Demonstrations and unique components of the 2.45 GHz beamed power transmission technology are reviewed. Attention is given to the following projects and developments: a microwave power helicopter demonstration; certification of overall dc to dc efficiency of 54 percent in the laboratory; demonstration of the transmission of significant amounts of power over significant distances; the solar power satellite and beamed power transmission; the Canadian SHARP program; space applications; an equatorially based beam power transmission system; an all-electronic LEO to GEO transportation system; and an orbiting industrial park system.
Parametric reentry study for elliptical parking and circular earth orbits
Comparison of two-impulse and three-impulse orbital transfer, using data from a 63-case numerical study. For each case investigated for which coplanarity of the regressing assembly parking ellipse was attained with the target asymptotic velocity vector, a two-impulse maneuver (or a one-impulse equivalent) was found for which the velocity expenditure was within 1% of a reference absolute minimum lower bound. Therefore, for the coplanar cases, use of a minimum delta-V three-impulse maneuver afforded scant improvement in velocity penalty. However, as the noncoplanarity of the parking ellipse and the target asymptotic velocity vector increased, there was a significant increase in the superiority of minimum delta-V three-impulse maneuvers for slowing the growth of velocity expenditure. It is concluded that a multiple-impulse maneuver should be contemplated if nonnominal launch conditions could occur.
NASA Marshall Space Flight Center (MSFC) is currently at work developing hardware and systems for the Ares I rocket that will send future astronauts into orbit. Built on cutting-edge launch technologies, evolved powerful Apollo and Space Shuttle propulsion elements, and decades of NASA spaceflight experience, Ares I is the essential core of a safe, reliable, cost-effective space transportation system -- one that will carry crewed missions back to the moon, on to Mars and out into the solar system. Ares I is an in-line, two-stage rocket configuration topped by the Orion crew vehicle and its launch abort system. In addition to the vehicle's primary mission -carrying crews of four to six astronauts to Earth orbit --Ares I may also use its 25-ton payload capacity to deliver resources and supplies to the International Space Station, or to "park" payloads in orbit for retrieval by other spacecraft bound for the moon or other destinations. Crew transportation to the International Space Station is planned to begin no later than 2014. The first lunar excursion is scheduled for the 2020 timeframe. This paper presents the challenges in designing the Ares I upper stage for reliability and safety while minimizing weight and maximizing performance.
A V-Slit sun sensor is body-mounted on a spin-stabilized spacecraft. During injection from a parking or transfer orbit to some final orbit, the spacecraft may not be dynamically balanced. This may result in wobble about the spacecraft spin axis as the spin axis may not be aligned with the spacecraft's axis of symmetry. While the widely used models in Spacecraft Attitude Determination and Control, edited by Wertz, correct for separation, elevation, and azimuthal mounting biases, spacecraft wobble is not taken into consideration. A geometric approach is used to develop a method for measurement of the sun angle which corrects for the magnitude and phase of spacecraft wobble. The algorithm was implemented using a set of standard mathematical routines for spherical geometry on a unit sphere.
High-order Taylor series expansions can be used to model nonlinear dynamics at the cost of integrating a large set of variational equations to obtain high-order state-transition tensors (STTs). This paper presents an innovative technique for approximating the high-order STTs that reduces significantly the computational cost by retaining only the dominant secular terms. We propagate the low-order partial derivatives of Kepler’s equation, which only requires the integration of six additional equations to extend an n-th order approximation to order (n + 1). The approximation stems from the Lindstedt-Poincare procedure and exploits the stability properties of orbital motion. Since the method makes no dynamical assumptions, it can accommodate any source of orbital perturbations. We show how the approximation of the second-order STT significantly increases the accuracy of the linear method for uncertainty propagation with only a small computational overhead. Finally, we derive a high-order approximate extended Kalman filter that implements the proposed approximation of the STT and improves the performance of linear filters. Examples of application with different perturbation sources include the heliocentric orbit of an asteroid, an orbiter around Europa, and an Earth-orbiting satellite.
The SARJ LITE (Lubrication Interval Test) test rig was built as a method to evaluate the performance of the grease repair on the Starboard SARJ of the International Space Station(ISS). The on-orbit SARJ was temporarily parked after receiving significant degradation on one of its race ring nitrided surfaces as a result of inadequate lubrication ( high dry contact friction) and unaccounted for roller traction kinematics. In a scaled down rig, flight like roller bearings were preloaded and cycled on a nitrided 15-5 race surface. Grease was added to the track and with instrumentation monitoring performance, trending data will be extracted and used to determine lubrication intervals for both Port and Starboard ISS SARJ's. The grease lubrication was found to be effective in eliminating the high friction that contributed to the on-orbit race degradation.