SEARCH · Engineering Papers
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.
Quote a phrase for an exact phrase match. Source license links do not imply unrestricted reuse.
Inexpensive latching relays incorporating one-second time delay prove effective for maintaining system power in place of computer facility automatic shutdown sensors that are activated by minute power surges or spikes in 60 Hz input signal.
The four Galilean satellites of Jupiter present a long-term collision hazard to an uncontrolled orbiting spacecraft that repeatedly enters the spatial region occupied by the satellites. Satellite close encounters and the likelihood of collision over a wide range of initial orbit conditions were analyzed. The effect of orbit inclination was of key interest. The scope of the analysis was restricted to orbital dynamic considerations alone, i.e. the question of biological contamination given the event of collision was not considered. A quarantine or orbiter lifetime of 50 years was assumed. This time period begins at spacecraft shutdown following completion of the mission objectives. A numerical approach was adopted wherein each initial orbit is propagated for 50 years, and satellite closest encounter distances recorded on every revolution. The computer program includes approximations of the three major perturbation effects on the long-term motion of the orbiter: (1) Jupiter oblateness, (2) solar gravity, and (3) satellite gravity.
The effects of afterburner light-off and shut-down transients on the compressor stability are investigated. The reported experimental results are based on detailed high response pressure and temperature measurements on the TF30-P-3 turbofan engine. The tests were performed in an altitude test chamber simulating high altitude engine operation. It is shown that during both types of transients, flow breaks down in the forward part of the fan bypass duct. At a sufficiently low engine inlet pressure this resulted in a compressor stall. Complete flow breakdown within the compressor was preceded by a rotating stall. At some locations in the compressor, rotating stall cells initially extended only through part of the blade span. For the shutdown transient the time between first and last detected occurrence of rotating stall is related to the flow Reynolds number. An attempt was made to deduce the number and speed of propagation of rotating stall cells.
The primary function of the landing gear is to dissipate touchdown energy in a controlled fashion that minimizes the landing shock for onboard components while maximizing ground clearance and the probability of landing stability over the range of possible surface and touchdown parameter variations. Many other mission requirements and constraints were translated into the evolving Viking landing gear design as they arose. Hence, design considerations included such factors as prelaunch heat sterilization and noncontamination of the Martian landing site, gear stowage and deployment, terminal descent engine shutdown initiation, structural load attenuation, hard/soft landing surface capability, reliability, weight, and post-landed stability.
In support of the ongoing NASA-European Space Agency (ESA) effort to understand and control possible interference between missions, testing was conducted at the Madrid Deep Space Station from July 1975 to February 1976 to characterize the effect on Viking 1975 telecommunication link performance of Geodetic Earth-Orbiting Satellite (GEOS) downlink signals. The prime use of the data was to develop a capability to predict GEOS interference effects for evaluation of Viking 1975 mission impacts and possible temporary GEOS shutdown. Also, the data would serve as a basis for assessment of the GEOS impact on missions other than Viking as well as for more general interference applications. Performances of the reference receiver, telemetry, and planetary ranging were measured in the presence of various types of GEOS-related interference, including an unmodulated GEOS carrier and simulation of the actual spectrum by an ESA-supplied GEOS suitcase model.
The feasibility of using methane, recovered from carbon dioxide by the Sabatier-Senderens reaction, was investigated as a fuel in a reaction control engine. A problem to be avoided is the emission of soot particles. It is thought that such particles would remain in the environment of the spacecraft and, thereby, adversely affect optical sightings and possibly have an effect upon communications. The initial studies were of a practical nature. The first was the influence of the spatial arrangement of the fuel and oxygen injectors on soot formation. In the second study, inhibition of soot formation was considered. Considering the given situation, it was impractical to use an additive. However, since methanol combustion does not produce soot, and methanol can be produced from methane in situ, the possibility that methanol could act as an inhibitor was studied. In the third study, since these are restartable engines, the effect of shutdown on the rapidity of soot formation was studied.
System architectures which incorporate fail operative flight guidance functions within a total integrated avionics complex are described. It is shown that the mixture of flight critical and nonflight critical functions within a common computer complex is an efficient solution to the integration of navigation, guidance, flight control, display, and flight management. Interfacing subsystems retain autonomous capability to avoid vulnerability to total avionics system shutdown as a result of only a few failures.
Power supply achieves efficiency rating of approximately 80 percent at rated maximum output current. Protection circuitry prevents logic/memory loads from losing stored data if supply protection circuit initiates power shut down; prevents damage caused by overtemperature condition; protects logic/memory load from undervoltage to overvoltage conditions. Warning signal indicates power shutdown.
The effect of ceramic coating of a JT8D combustor liner was investigated at simulated cruise and takeoff conditions with two fuels of widely different aromatic contents. Substantial decreases in maximum liner temperatures and flame radiation values were obtained with the ceramic-coated liner. Small reductions in exhaust gas smoke concentrations were observed with the ceramic-coated liner. Other performance parameters such as combustion efficiency and emissions of unburned hydrocarbons, CO, and NOx were not affected significantly. No deterioration of the ceramic coating was observed after about 6 hours of cyclic operation including several startups and shutdowns.
The Wind Turbine demonstrates the capability of automatic unattended operation, including startup, achieving synchronism, and shutdown as dictated by wind conditions. During the course of these operations, a wealth of engineering data was generated. Some of the data which is associated with rotor and machine dynamics problems encountered, and the machine modifications incorporated as a solution are presented. These include high blade loads due to tower shadow, excessive nacelle yawing motion, and power oscillations. The results of efforts to correlate measured wind velocity with power output and wind turbine loads are also discussed.
The results of data obtained 80 days preceding the solar eclipse shutdown of the CTS 11.7 GHz righthand circularly polarized beacon transmitter are presented. Attenuation and cross polarization isolation were measured. It was determined that depolarization presents a serious limitation to satellite system reliability when frequency reuse by polarization diversity is employed. A 27 db isolation margin would reduce reliability below 99.95%. For the same percentage the required fade margin was below 3 db.
Development of a recuperator for a 2.0 kW closed Brayton space power system is described. The plate-fin heat exchanger is fabricated entirely from Hastelloy X and is designed for 10 years continuous operation at 1000 K (1300 F) with a Xenon-helium working fluid. Special design provisions assure uniform flow distribution, crucial for meeting 0.975 temperature effectiveness. Low-cycle fatigue, resulting from repeated startup and shutdown cycles, was identified as the most critical structural design problem. It is predicted that the unit has a minimum fatigue life of 220 cycles. This is in excess of the BIPS requirement of 100 cycles. Heat transfer performance and thermal cycle testing with air, using a prototype unit, verified that all design objectives can be met.
The use of a recoverable liquid rocket booster (LRB) system to replace the existing solid rocket booster (SRB) system for the shuttle was studied. Historical weight estimating relationships were developed for the LRB using Saturn technology and modified as required. Mission performance was computed using February 1975 shuttle configuration groundrules to allow reasonable comparison of the existing shuttle with the study designs. The launch trajectory was constrained to pass through both the RTLS/AOA and main engine cut off points of the shuttle reference mission 1. Performance analysis is based on a point design trajectory model which optimizes initial tilt rate and exoatmospheric pitch profile. A gravity turn was employed during the boost phase in place of the shuttle angle of attack profile. Engine throttling add/or shutdown was used to constrain dynamic pressure and/or longitudinal acceleration where necessary. Four basic configurations were investigated: a parallel burn vehicle with an F-1 engine powered LRB; a parallel burn vehicle with a high pressure engine powered LRB; a series burn vehicle with a high pressure engine powered LRB. The relative sizes of the LRB and the ET are optimized to minimize GLOW in most cases.
Historical weight estimating relationships were developed for the liquid rocket booster (LRB) using Saturn technology, and modified as required to support the EDIN05 study. Mission performance was computed using February 1975 shuttle configuration groundrules to allow reasonable comparison of the existing shuttle with the EDIN05 designs. The launch trajectory was constrained to pass through both the RTLS/AOA and main engine cut-off points. Performance analysis was based on a point design trajectory model which optimized initial tilt rate and exo-atmospheric pitch profile. A gravity turn was employed during the boost phase in place of the shuttle angle-of-attack profile. Engine throttling add/or shutdown was used to constrain dynamic pressure and/or longitudinal acceleration where necessary.
Long term tests were performed on a 30 cm Hg bombardment thruster and a power processing unit to determine lifetime characteristics. The thruster performance data and other operational characteristics taken at various times during the test segment are presented and evaluated with the life limiting mechanisms: discharge chamber erosion, deposition and spalling, external erosion, cathode degradation, and propellant isolator leakage. The control algorithms for thruster start up, steady state operation, throttle, detection and correction of off normal conditions, and shutdown are discussed.
Filament of high-power radio transmitter is brought to operating temperature automatically. Pushbotton reduces operator's role to one-step command and is compatible with various forms of computer control. Filiament shutdown is initiated by "down" command from operator, failure of cooling systems, or power failure for more than few seconds.
Report describes system which meets 64 percent of hot water requirements of 120 room motel. Key system components include 1,000 square foot, roof-mounted collector array, 1,000 gallon storage tank, tube-in-shell heat exchanger, and three domestic hot-water tanks. Report contains calibration instructions for differential temperature controllers, shutdown procedures, and operation guidelines, performance analysis, and manufacturers' maintenance literature.
The preprototype water quality monitor (WQM) subsystem was designed based on a breadboard monitor for pH, specific conductance, and total organic carbon (TOC). The breadboard equipment demonstrated the feasibility of continuous on-line analysis of potable water for a spacecraft. The WQM subsystem incorporated these breadboard features and, in addition, measures ammonia and includes a failure detection system. The sample, reagent, and standard solutions are delivered to the WQM sensing manifold where chemical operations and measurements are performed using flow through sensors for conductance, pH, TOC, and NH3. Fault monitoring flow detection is also accomplished in this manifold assembly. The WQM is designed to operate automatically using a hardwired electronic controller. In addition, automatic shutdown is incorporated which is keyed to four flow sensors strategically located within the fluid system.