Solid rocket propulsion
The solid rocket motor failure database; current programs; and shortfalls of solid rocket propulsion are examined. This presentation is represented by viewgraphs.
Nichols, Ronald L.↗
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A review of liquid rocket propulsion programs in Japan
An assessment of Japan's current capabilities in the areas of space and transatmospheric propulsion is presented. The primary focus is upon Japan's programs in liquid rocket propulsion and in space plane and related transatmospheric areas. Brief reference is also made to their solid rocket programs, as well as to their supersonic air breathing propulsion efforts that are just getting underway.
Merkle, Charles L.↗
Experimental and analytical comparison of flowfields in a 110 N (25 lbf) H2/O2 rocket
A gaseous hydrogen/gaseous oxygen 110 N (25 lbf) rocket was examined through the RPLUS code using the full Navier-Stokes equations with finite rate chemistry. Performance tests were conducted on the rocket in an altitude test facility. Preliminary parametric analyses were performed for a range of mixture ratios and fuel film cooling pcts. It is shown that the computed values of specific impulse and characteristic exhaust velocity follow the trend of the experimental data. Specific impulse computed by the code is lower than the comparable test values by about two to three percent. The computed characteristic exhaust velocity values are lower than the comparable test values by three to four pct. Thrust coefficients computed by the code are found to be within two pct. of the measured values. It is concluded that the discrepancy between computed and experimental performance values could not be attributed to experimental uncertainty.
Reed, Brian D.↗
Reusable rocket engine intelligent control system framework design, phase 2
Elements of an advanced functional framework for reusable rocket engine propulsion system control are presented for the Space Shuttle Main Engine (SSME) demonstration case. Functional elements of the baseline functional framework are defined in detail. The SSME failure modes are evaluated and specific failure modes identified for inclusion in the advanced functional framework diagnostic system. Active control of the SSME start transient is investigated, leading to the identification of a promising approach to mitigating start transient excursions. Key elements of the functional framework are simulated and demonstration cases are provided. Finally, the advanced function framework for control of reusable rocket engines is presented.
Nemeth, ED↗
Discussions on a long gap discharge to an EHV transmission tower by a rocket triggered lightning experiment
The triggered lightning experiments using a rocket have been carried out on a winter mountain in Japan since 1986. For the four years from 1986 to 1989, 39 rockets were launched and 19 of them triggered lightning strikes. The emphasis here is on the methodology for triggering lightning to the transmission system. Completed experiments are discussed. The failure of lightning protection and the striking distance are noted.
Nakamura, Koichi↗
Sounding rocket lessons learned
Programmatic, applicatory, developmental, and operational aspects of sounding rocket utilization for materials processing studies are discussed. Lessons learned through the experience of 10 sounding rocket missions are described. Particular attention is given to missions from the SPAR, Consort, and Joust programs. Successful experiments on Consort include the study of polymer membranes and resins, biological processes, demixing of immiscible liquids, and electrodeposition.
Wessling, Francis C.↗
Intense turbulence in the polar mesosphere - Rocket and radar measurements
Two radars respectively operating at 53.5 MHz and 224 MHz have been used to simultaneously measure the polar mesospheric summer echoes (PMSEs) during the launch of a sounding rocket designed to measure the fluctuations in the electron density in the same height range. Some features of the data obtained are consistent with nearly isotropic, intense, meter-scale turbulence on the day of the observations. A dissipation rate is derived from the spectral width of the 53.5 MHz backscatter signal and from the rocket electron density fluctuation data; the diffusive subrange occurs at a wavelength which is a factor of 10 greater than the viscous subrange. This corresponds to a Schmidt number of about 100.
Kelley, M. C.↗
Spectrally resolved Rayleigh scattering diagnostic for hydrogen-oxygen rocket plume studies
A Rayleigh scattering diagnostic has been developed to measure gas density, temperature, and velocity in the exhaust plume of 100 N thrust class hydrogen-oxygen rockets. The spectrum of argon-ion laser light scattered by the gas molecules in the plume (predominantly water vapor) is measured with a scanning Fabry-Perot interferometer. The gas density is determined from the total scattered power, the gas temperature from the spectral width, and the velocity from the shift in the peak of the spectrum from the frequency of the incident laser light. The diagnostic has been demonstrated in a rocket test cell and a discussion of results is given.
Seasholtz, R. G.↗
Electric field measurements during supercharging events on the MAIMIK rocket experiment
Data are presented on the evolution of electric field structures near the MAIMIK mother-daughter tethered rocket during active electron emissions from the rocket, which consisted of two separately instrumented payload sections which were detached in-flight. The results of the E field data analysis show that the temporal behavior of the fields was consistent with ion responses to the rapid charging of the active payload. A qualitative model of the interaction between the active payload and the ambient plasma background is developed to explain the electric field data, including the apparent causal relationship between the radial and tangential impulse components.
Denig, W. F.↗
Status report on a real time Engine Diagnostics Console for rocket engine exhaust plume monitoring
This paper describes the work done on the Engine Diagnostics Console during the past year of development at Stennis Space Center. The Engine Diagnostics Console (EDC) is a hardware and software package which provides near real time monitoring of rocket engine exhaust plume emissions during ground testing. The long range goal of the EDC development program is to develop an instrument that can detect engine degradation leading to catastrophic failure, and respond by taking preventative measures. The immediate goal for the past year's effort is the ability to process spectral data, taken from a rocket engine's exhaust plume, and to identify in an automated and high speed manner, the elemental species and multielemental materials that are present in the exhaust plume.
Bircher, F. E.↗
The use of programmable logic controllers (PLC) for rocket engine component testing
Application of PLCs to the rocket engine component testing at a new Stennis Space Center Component Test Facility is suggested as an alternative to dedicated specialized computers. The PLC systems are characterized by rugged design, intuitive software, fault tolerance, flexibility, multiple end device options, networking capability, and built-in diagnostics. A distributed PLC-based system is projected to be used for testing LH2/LOx turbopumps required for the ALS/NLS rocket engines.
Nail, William↗
Modeling flow at the nozzle of a solid rocket motor
The mechanical behavior of a rocket motor internal flow field results in a system of nonlinear partial differential equations which can be solved numerically. The accuracy and the convergence of the solution of the system of equations depends largely on how precisely the sharp gradients can be resolved. An adaptive grid generation scheme is incorporated into the computer algorithm to enhance the capability of numerical modeling. With this scheme, the grid is refined as the solution evolves. This scheme significantly improves the methodology of solving flow problems in rocket nozzle by putting the refinement part of grid generation into the computer algorithm.
Chow, Alan S.↗
Ascent performance issues of a vertical-takeoff rocket launch vehicle
Advanced manned launch systems studies under way at the NASA Langley Research Center are part of a broader effort that is examining options for the next manned space transportation system to be developed by the United States. One promising concept that uses near-term technologies is a fully reusable, two-stage vertical-takeoff rocket vehicle. This vehicle features parallel thrusting of the booster and orbiter with the booster cross-feeding the propellant to the orbiter until staging. In addition, after staging, the booster glides back unpowered to the launch site. This study concentrated on two issues that could affect the ascent performance of this vehicle. The first is the large gimbal angle range required for pitch trim until staging because of the propellant cross-feed. Results from this analysis show that if control is provided by gimballing of the rocket engines, they must gimbal greater than 20 deg, which is excessive when compared with current vehicles. However, this analysis also showed that this limit could be reduced to 10 deg if gimballing were augmented by throttling the booster engines. The second issue is the potential influence of off-nominal atmospheric conditions (density and winds) on the ascent performance. This study showed that a robust guidance algorithm could be developed that would insure accurate insertion, without prelaunch atmospheric knowledge.
Powell, Richard W.↗
In-situ propellant rocket engines for Mars missions ascent vehicle
When contemplating the human exploration of Mars, many scenarios using various propulsion systems have been considered. One propulsion option among them is a vehicle stage with multiple, pump fed rocket engines capable of operating on propellants available on Mars. This reduces the earth launch mass requirements, resulting in economic and payload benefits. No plentiful sources of hydrogen on Mars have been identified on the surface of Mars, so most commonly used high performance liquid fuels, such as hydrogen and hydrocarbons, can be eliminated as possible in situ propellants. But 95 pct of the Martian atmosphere consists of carbon dioxide, which can be converted into carbon monoxide and oxygen. The carbon monoxide oxygen propellant combination is a candidate for a Martian in situ propellant rocket engine. The feasibility is analyzed of a pump fed engine cycle using the propellant combination of carbon monoxide and oxygen.
Roncace, Elizabeth A.↗
The rationale/benefits of nuclear thermal rocket propulsion for NASA's lunar space transportation system
Two nuclear thermal rocket (NTR) technology options are examined - one derived from the graphite-moderated reactor concept developed by NASA and the AEC under the Rover/NERVA (Nuclear Engine for Rocket Vehicle Application) programs, and a second concept, the Particle Bed Reactor. The paper also summarizes NASA's lunar outpost scenario, compares relative performance provided by different lunar space transportation system concepts, and discusses important operational issues (e.g., reusability, engine 'end-of-life' disposal, etc.) associated with using this important propulsion technology.
Borowski, Stanley K.↗
Experimental and analytical comparison of flowfields in a 110 N (25 Lbf) H2/O2 rocket
A gaseous hydrogen/gaseous oxygen 110 N (25 lbf) rocket has been examined through the RPLUS code using the full Navier-Stokes equations with finite-rate chemistry. Performance tests were conducted on the rocket in an altitude test facility. Preliminary parametric analyses have been performed for a range of mixture ratios and fuel film cooling percentages. It is shown that the computed values of specific impulse and characteristic exhaust velocity follow the trend of the experimental data. Specific impulse computed by the code is lower than the comparable test values by about two to three percent. The computed characteristic exhaust velocity values are lower than the comparable test values by three to four percent. Thrust coefficients computed by the code are found to be within two percent of the measured values. It is concluded that the discrepancy between computed and experimental performance values could not be attributed to experimental uncertainty.
Reed, Brian D.↗
Rocket measurements of relativistic electrons - New features in fluxes, spectra and pitch angle distributions
Novel features are presented of precipitating relativistic electron fluxes measured on a spinning sounding rocket payload at midday between altitudes of 70 and 130 km in the auroral region. The sounding rocket was launched during a relativistic electron enhancement event of modest intensity. Electron fluxes were measured for a total of about 210 seconds at energies from 0.1 to 3.8 MeV, while pitch angle was sampled from 0 to 90 deg every spin cycle. Electrons with energies below about 0.2 MeV showed isotropic pitch angle distributions during most of the first 90 sec of data, while at higher energies the electrons had the highest fluxes near the mirroring angle (90 deg); when they occurred, the noted downward bursts were seen at all energies. The low-altitude fluxes are compared with those measured at geostationary orbit, and it is found that the low-altitude fluxes are much higher than expected from a simple mapping of a pancake distribution at high altitudes.
Herrero, F. A.↗
Computational fluid dynamics studies of nuclear rocket performance
A CFD analysis of a low pressure nuclear rocket concept is presented with the use of an advanced chemical kinetics, Navier-Stokes code. The computations describe the flow field in detail,including gas dynamic, thermodynamic and chemical properties, as well as global performance quantities such as specific impulse. Computational studies of several rocket nozzle shapes are conducted in an attempt to maximize hydrogen recombination. These Navier-Stokes calculations, which include real gas and viscous effects, predict lower performance values than have been reported heretofore.
Stubbs, Robert M.↗