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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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Water vapor is absorbed in hygroscopic electrolyte, and oxygen generated by absorbed water electrolysis at anode is added simultaneously to air stream. Cell applications include on-board aircraft oxygen systems, portable oxygen generators, oxygen concentration requirements, and commercial air conditioning and dehumidifying systems.
Successful planetary exploration missions have been completed by unmanned Mariner spacecraft, which were designed at the Jet Propulsion Laboratory. Mariner II completed a 109-day flyby mission to Venus on December 14, 1962. This mission was followed by the flyby of Mars and return of 21 historic photographs of its surface by Mariner IV in July 1965. Further successful missions followed: Mariner V to Venus in 1967, Mariners VI and VII to Mars in 1969. In all of these missions, a Central Computer and Sequencer onboard the spacecraft was used to provide on-board control functions, although complete ground backup was also available.
Handling of waste products on spacecraft to include identification, potential for on-board utilization, pretreatment requirements for storage, and disposal concepts - Vol.1
Trapped protons east-west asymmetry observations during Gemini 4 flight, using on-board high sensitivity plastic-scintillation spectrometer
Bull frog activity at rest and response to centripetal acceleration by on-board centrifuge in vestibular space experiment OFO-A
The existing TNT equivalency criterion for LO2/LH2 propellant is reevaluated. It addresses the static, on-pad phase of the space shuttle launch operations and was performed to determine whether the use of a TNT equivalency criterion lower than that presently used (60%) could be substantiated. The large quantity of propellant on-board the space shuttle, 4 million pounds, was considered of prime importance to the study. A qualitative failure analysis of the space shuttle (EOS) on the launch pad was made because it was concluded that available test data on the explosive yield of LO2/LH2 propellant was insufficient to support a reduction in the present TNT equivalency value, considering the large quantity of propellant used in the space shuttle. The failure analysis had two objectives. The first was to determine whether a failure resulting in the total release of propellant could occur. The second was to determine whether, if such a failure did occur, ignition could be delayed long enough to allow the degree of propellant mixing required to produce an explosion of 60% TNT equivalency since the explosive yield of this propellant is directly related to the quantities of LH2 and LO2 mixed at the time of the explosion.
Investigation of problems related to control of a mobile planetary vehicle according to a systematic plan for the exploration of Mars has been undertaken. Problem areas receiving attention include: (1) overall systems analysis; (2) vehicle configuration and dynamics; (3) toroidal wheel design and evaluation; (4) on-board navigation systems; (5) satellite-vehicle navigation systems; (6) obstacle detection systems; (7) terrain sensing, interpretation and modeling; (8) computer simulation of terrain sensor-path selection systems; and (9) chromatographic systems design concept studies. The specific tasks which have been undertaken are defined and the progress which has been achieved during the period July 1, 1971 to December 31, 1971 is summarized.
A preliminary study has been made of the navigation performance which might be achieved for the high cross-range space shuttle orbiter during final approach and landing by using an optimally augmented inertial navigation system. Computed navigation accuracies are presented for an on-board inertial navigation system augmented (by means of an optimal filter algorithm) with data from two different ground navigation aids; a precision ranging system and a microwave scanning beam landing guidance system. These results show that augmentation with either type of ground navigation aid is capable of providing a navigation performance at touchdown which should be adequate for the space shuttle. In addition, adequate navigation performance for space shuttle landing is obtainable from the precision ranging system even with a complete dropout of precision range measurements as much as 100 seconds before touchdown.
An estimation of the thrust magnitude of the Space Electric Rocket Test (SERT) engine is obtained. This estimation is based on changes in the SERT-2 spacecraft orbit as determined from tracking data by using a revised definitive orbit determination system which includes the effect of thrust upon the motion of a spacecraft in general and the SERT-2 satellite in particular. The results of these computations are found to be in good agreement with values of acceleration as determined by an on-board accelerometer. A discussion of the mathematical model used to obtain the numerical results is also presented.
A brief summary of the mechanical properties of Jupiter is presented along with an organizational outline of the entire JOSE program. Other aspects of the program described include: spacecraft design, mission trajectories, altitude control, propulsion subsystem, on-board power supply, spacecraft structures and environmental design considerations, and telemetry.
A study which illustrated the feasibility of implementing an on-board aircraft navigation system based on using the ATS-1 and ATS-3 satellites, the modified Omega Position Location Equipment (OPLE) Control Center, and a suitable aircraft terminal was conducted. The report provides: (1) a consideration of the problems of satellite navigation and an objective definition of the optimum system under the constraints of its specified components, (2) a description of the necessary modifications to the OPLE Control Center, the design of an aircraft terminal, and the design of ground reference terminals, and (3) an outline of an experiment plan and an estimate of the cost to be expected in conducting the program.
During the re-entry phase of a manned satellite, some equipment for continuous on-board indication of position will be required. Since radio and radar may be useless during part of the re-entry, inertial guidance equipment may be required. Such equipment, however, has an inherent instability in the computation of altitude. The present study of an inertial guidance system shows that for reasonable values of initial-condition errors and equipment biases. the resultant position indication errors will not become excessive unless the re-entry maneuver time is greater than 45 minutes to an hour. Further, the position indication error caused by accelerometer uncertainties can be reduced by switching off the accelerometers until their output becomes significantly greater than their uncertainty.
A series of balloon-borne experiments has been initiated at the Oak Ridge National Laboratory (ORNL) in cooperation with the Space Sciences Laboratory of the NASA G. C. Marshall Space Flight Center to determine the feasibility of using large (greater than 20 cc) lithium-drifted, germanium diodes to investigate the spectrum of atmospheric gamma rays (60 keV < E < 1.5 MeV) at altitudes corresponding to a few grams per cm (sup 2) residual atmosphere. The results of these measurements will also provide a basis for designing an appropriate shield for a multi-diode, highly directional gamma-ray spectrometer for use in astrophysical measurements. Two flights were accomplished during 1967 using the same two diodes and basic instrument package. Both flights were launched at the NCAR Scientific Balloon Flight Base, Palestine, Texas, and attained a float altitude of 117, 000 ft. The measured atmospheric gamma-ray spectrum shows at least one distinct line at 511 keV (annihilation radiation) superimposed on a continuous distribution of gamma rays attributable to both energy-degraded gamma rays and brems-strahlung photons. A second diode, encased in a passive shield of plastic and lead, shows the expected lines resulting from both neutron inelastic scattering and capture in the germanium. Data acquisition was accomplished on an event-by-event basis through the use of 512-channel, on-board ADC and word-generator circuits and a ground-based, on-line telemetry decoder. The decoder makes it possible to store the data in the memory of a modified pulse-height analyzer simultaneously with storage on magnetic tape. This provides a real-time visual observation of the data as it is accumulated and greatly facilitates preflight calibrations.
Analytical techniques are presented that permit the calculation of heat-transfer rates with various thermal-protection systems for liquid-cryogenic-propellant tanks subjected to on-board, solar, and planetary heat fluxes . The effectiveness of these protection systems in reducing propellant heating is shown both for ideal heat-transfer models and for a simplified hydrogen-oxygen terminal stage used for typical Mars missions.
Statistical filter theory is employed to develop a method for determining the best possible estimate of the position and velocity of a space vehicle in the midcourse phase of flight. Results of a computer simulation are given to illustrate the performance attainable. An on-board system is visualized in which the source of information is an arbitrary sequence of observations of space angles, corrupted by measurement errors. The scheme is in effect a dynamical time-varying filter, implemented by a digital computer, which processes the incoming data to compute an up-to-date optimal estimate of position and velocity.
This paper addresses the problem of building trust in online predictions of the remaining available flying time for two different electric Unmanned Aerial Vehicles (eUAVs) powered by lithium-ion-polymer batteries. Flight tests for various automation research missions for the two vehicles were monitored using two on-board battery health management applications to make predictions of the remaining flying time (RFT) for each eUAV and to predict the state of the battery. Playback of the voltage, current and temperature profiles of the battery discharge were used to assess the accuracy of the estimation of the voltage and the charge states of the models as well as the estimate of the RFT. The reference ground truth values were the observed landing time and the measured battery pack resting pack voltage 20 minutes after the flight. The predicted RFT, state of charge (SoC), and state of energy (SoE) were compared with the observed results. Noise values of one standard deviation from the mean values of the internal charge states of the battery model during a reference run were used to vary the states during simulation. One application used an equivalent circuit model of the electrical dynamics of the battery pack, and the other application used a reduced-order electrochemistry model. The variation of the model state components was compared to the variation in the estimate of the RFT and the variation in the SoE to estimate a confidence factor. Variation in the estimates caused by factors affecting the off-line laboratory parameter identification experiments is considered. Variation in the estimates due to environmental factors are discussed.
NASA aeronautics goals include pioneering new technology to increase efficiency and reduce emissions from air travel. The feasibility of using of on-board cryogenic fuels systems in combination with fuel cells, high efficiency engines, and electric motors for high efficiency commercial transport aircraft is being examined at NASA in collaborations with external partners. Growing emphasis on next generation air transportation resulted in the development of strategic investigations into cryogenic fuels aviation, including roadmaps for maturing new technologies, needs to improve testing and manufacturing capabilities, and the requirements for establishing regulatory and certification standards. A NASA cross-organizational multidisciplinary project team is exploring new aircraft architectures, enabling materials, and the engineering challenges associated with cryogenic fuels for aircraft. The engineering challenges include high performance cryogenic, thermal management, and power/propulsion components and subsystems to tackle the challenging problem of development of commercially viable aircraft that would radically transform air transportation.
Advanced electrified aircraft propulsion (EAP) concepts with integrated power, propulsion, and thermal systems require the development of equally sophisticated controllers to fly safe and efficient missions. Hybrid electric aircraft utilize electric machines mechanically coupled to the engine shafts to extract and insert power for a variety of purposes. Additionally, electric machines may drive propulsive fans pulling from a combination of on-board energy storage devices and engine extracted power. NASA’s investment in hybrid electric aircraft hardware-in-the-loop testing enables controls research on representative engine models using a novel emulation and scaling methodology. High-voltage, high-power electrical powertrain presents several technical risks at altitude. High-power requires high efficiency to minimize losses. Superconducting electric machines and power distribution research at NASA has identified key challenges with potential controls solutions. These risks necessitate the development of controllers robust to model uncertainty, disturbances, and fault conditions. System health management and fault detection schemes play an important role in a multi-layer controls approach that utilizes a supervisor which oversees inner loops for the highly coupled power, propulsion, and thermal systems. Such a supervisory controller enables coordinates energy transfer between engines, energy storage devices, electric machines, propulsors, and heat exchangers. Existing methods have shown an improvement in engine operability using the hybrid electric powertrain.