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The Method of Integral Relations is extended to apply to the calculation of Unsteady Flow in a Laminar Boundary Layer. The effectiveness of the method is demonstrated in an application to uniform flow past a circular cylinder, resulting from an impulsive start. Both first and second approximations are worked out, the latter showing good agreement with results of previous calculations.
A simple model of tidal encounters has been applied to the problem of an open cluster in a clumpy molecular cloud. The parameters of the clumps are taken from the Blitz, Stark, and Long (1988) catalog of clumps in the Rosette molecular cloud. Encounters are modeled as impulsive, rectilinear collisions between Plummer spheres, but the tidal approximation is not invoked. Mass and binding energy changes during an encounter are computed by considering the velocity impulses given to individual stars in a random realization of a Plummer sphere. Mean rates of mass and binding energy loss are then computed by integrating over many encounters. Self-similar evolutionary calculations using these rates indicate that the disruption process is most sensitive to the cluster radius and relatively insensitive to cluster mass. The calculations indicate that clusters which are born in a cloud similar to the Rosette with a cluster radius greater than about 2.5 pc will not survive long enough to leave the cloud. The majority of clusters, however, have smaller radii and will survive the passage through their parent cloud.
Analyses were conducted which indicate that sub kW-class ion thrusters may provide performance benefits for near-Earth space commercial and science missions. Small spacecraft applications with masses ranging from 50 to 500 kg and power levels less than 0.5 kW were considered. To demonstrate the efficacy of propulsion systems of this class, two potential missions were chosen as examples; a geosynchronous north-south station keeping application, and an Earth orbit magnetospheric mapping satellite constellation. Xenon ion propulsion system solutions using small thrusters were evaluated for these missions. A payload mass increase of more than 15% is provided by a 300-W ion system for the north-south station keeping mission. A launch vehicle reduction from four to one results from using the ion thruster for the magnetospheric mapping mission. Typical projected thruster performance over the input power envelope of 100-300 W range from approximately 40% to 54% efficiency and approximately 2000 to 3000 seconds specific impulse. Thruster technologies required to achieve the mission-required performance and lifetime are identified.
NTP development is currently supported by the NASA program office "Advanced Exploration Systems". The concept is a main propulsion option being considered for human missions to Mars in the 2030's. Major NTP development took place in the 1960's and 1970's under the Rover/NERVA program. The technology had matured to TRL 6 and was preparing to go to TRL 7 with a prototype flight engine before the program was cancelled. Over the last 40 years, a variety of continuations started, but only lasted a few years each. The Rover/NERVA infrastructure is almost all gone. The only remains are a few pieces of hardware, final reports and a few who worked the Rover/NERVA. Two types of nuclear fuel are being investigated to meet the current engine design specific impulse of 900 seconds compared to approximately 850 seconds demonstrated during Rover/NERVA. One is a continuation of composite fuel with new coatings to better control mid-band corrosion. The other type is a CERMET fuel made of Tungsten and UO2. Both fuels are being made from Rover/NERVA lessons learned, but with slightly different recipes to increase fuel endurance at higher operating temperatures. The technology readiness level (TRL) of these current modified reactor fuels is approximately TRL 3. To keep the development cost low and help mature the TRL level past 4 quickly, a few special non-nuclear test facilities have been made to test surrogate fuel, with depleted uranium, as coupons and full length elements. Both facilities utilize inductive heating and are licensed to handle depleted uranium. TRL 5 requires exposing the fuel to a nuclear environment and TRL 6 requires a prototype ground or flight engine system test. Currently, three different NTP ground test facility options are being investigated: exhaust scrubber, bore hole, and total exhaust containment. In parallel, a prototype flight demonstration test is also being studied. The first human mission to Mars in the 2030's is currently 2033. For an advanced propulsion concept to be seriously considered for use, the engine development plans need to show it is feasible and affordable to reach TRL 8 by 2027 and can be qualified for human mission use.
Presented are several methods to study a system of tiles bonded elastically to a finite flexible plate responding to an impulse. First, a one-dimensional (1-D) approximation to the system is constructed using Timoshenko beam theory and transfer matrices. This simulates a strip one tile wide. Modal analysis determines the transient response to an impulse of short duration. Also presented are parametric studies of the effect of bond stiffness and the effect of stiffening from inplane stresses produced by large out-of-plane deflections. A two-dimensional (2-D) model is constructed from the eigenfunctions of the 1-D model in a Galerkin method using pairwise products of the 1-D modes as trial functions. Comparison of the two models establishes a factor applied to peak impulse pressure for results to agree.
Despite the evidence from past earthquakes and several numerical investigations demonstrating the detrimental impact of vertical ground motions (VGMs) on the integrity of bridge structures, incorporating their effects into seismic assessment and design procedures has traditionally been given limited consideration. Current codes utilize rather simplistic approaches to account for the concurrent effects of vertical and horizontal motions in structural performance evaluations, potentially leading to unconservative estimates of structural demands. This paper reviews the main features of VGMs and their effect on the seismic response of bridges. The methods and empirical models available to estimate vertical motions for design purposes are discussed, and research gaps and related research needs are identified. Finally, the emerging role of physics-based ground-motion simulations, as well as their limitations, in supporting future research and informing the development of simplified design procedures is examined. The main areas of interest for future research are identified in the need to carry out systematic sensitivity studies to gain insight into the main earthquake parameters that influence key VGMs features, understand the influence of soil nonlinearities on VGMs amplitude and frequency content, inform the development of empirical models that cover a range of site conditions and source-to-site distances where current models are poorly constrained, generate arrays of motions to update coherency models to properly inform the analysis of distributed infrastructure, investigate the impulsive character of VGMs, and assess the approximations made in estimating VGMs with 1D site response analyses. Furthermore, specific focus is laid on large-magnitude earthquakes in the near-field.
The results are reported of a program to develop special instrumentation systems and engine hardware, conduct tests using LOX/GH2 propellants wherein radial mixtures ratio stratification was controlled, and subsequently compare the results of four selected tests with the predictions of the JANNAF performance-prediction computer programs. During the experiments, the overall propellant mixture ratio was varied from 4.4 to 6.6, while the mixture ratios in the core and outer zone were varied from 5.7 to 8.8 and from 3.7 to 7.2, respectively. A nominal 10 percent of the total fuel flow was used as boundary layer collant in a majority of the firings. Nominal chamber pressure was either 225 or 250 psia, with nozzle expansion ratios of either 25:1 or 4:1. Measurements of the axial chamber pressure and wall heat flux profiles, together with samples of the exhaust gas, were obtained. The corrected experimental specific impulse and characteristic exhaust velocity efficiencies were approximately 97.5 and 98.5 percent, respectively.
A detailed study was made of the lifetimes and evolution of fibrils in McMath 12417, using high resolution filtergrams in H alpha and CaII K made at Big Bear Solar Observatory. It was found that the lifetime of a fibril increases monotonically with its length. This relationship, together with the form of the variation of fibril lengths as a function of time, suggests that fibrils result from material being impulsively injected into magnetic field lines at approximately 30 km/sec, and returning under gravity. The lifetimes and apparent lengths of fibrils are then a function of the inclination of the field lines only. A study of wavelength scans through the H alpha line confirms that the apparent expansion and contraction of fibrils represents true mass motion.
A detailed study has been made of the lifetimes and evolution of fibrils in McMath 12417, using high resolution filtergrams in H-alpha and Ca II K. It was found that when viewed near disk center, the lifetime of a fibril is a monotonically increasing function of its maximum apparent length. This relationship, together with the form of the variation of fibril lengths as a function of time, suggests that fibrils result from material being impulsively injected into magnetic field lines at approximately 30 km/sec, and returning under gravity. The lifetimes and apparent lengths of fibrils are then a function of the inclination of the field lines only. A study of wavelength scans through the H-alpha line confirms that the apparent extension and retraction of fibrils represents true mass motion.
Laser thermal propulsion (LTP) is studied for the case in which laser power is absorbed by a small very high-temperature plasma (about 20,000 K) and transferred to the remainder of the pure hydrogen propellant by radiation and mixing. This concept could lead to the realization of a lightweight orbital transfer vehicle propulsion system having a specific impulse in the range 1000-2000 s. Approximately 12 percent of the input power may be radiated to the thruster walls, and 15 percent of the total propellant flow must be heated to 20,000 K to provide a bulk temperature of 5000 K prior to expansion. Three principal research issues identified are: (1) conditions for hydrogen plasma ignition, (2) control of the plasma position within the laser beam, plasma stability, and plasma absorption efficiency, and (3) characterization of the mixing of the plasma and buffer flows.
Conventional observational data regarding solar millisecond spikes are compared with data gathered simultaneously in the hard X-ray band by means of a statistical analysis. The analysis considers the association rate, correlation degree, and relative time delays between hard X-ray emissions (in the 25-438 keV range) and radio-spike events. About 95 percent of the radio-spike bursts occur during impulsive hard X-ray bursts, and approximately 43 percent of the compared events are characterized by hard X-ray time profiles that mimic the concentration of simultaneous radio spikes. The delay of the radio emission with respect to the hard X-ray bursts puts some constraints on the acceleration and propagation of particles. The time delays and the quantization into discrete radio events are theorized to be caused by the operation of the accelerator.
An analysis of gas turbine engines using water and oxygen injection to enhance performance by increasing Mach number capability and by increasing thrust is described. The liquids are injected, either separately or together, into the subsonic diffuser ahead of the engine compressor. A turbojet engine and a mixed-flow turbofan engine (MFTF) are examined, and in pursuit of maximum thrust, both engines are fitted with afterburners. The results indicate that water injection alone can extend the performance envelope of both engine types by one and one-half Mach numbers at which point water-air ratios reach 17 or 18 percent and liquid specific impulse is reduced to some 390 to 470 seconds, a level about equal to the impulse of a high energy rocket engine. The envelope can be further extended, but only with increasing sacrifices in liquid specific impulse. Oxygen-airflow ratios as high as 15 percent were investigated for increasing thrust. Using 15 percent oxygen in combination with water injection at high supersonic Mach numbers resulted in thrust augmentation as high as 76 percent without any significant decrease in liquid specific impulse. The stoichiometric afterburner exit temperature increased with increasing oxygen flow, reaching 4822 deg R in the turbojet engine at a Mach number of 3.5. At the transonic Mach number of 0.95 where no water injection is needed, an oxygen-air ratio of 15 percent increased thrust by some 55 percent in both engines, along with a decrease in liquid specific impulse of 62 percent. Afterburner temperature was approximately 4700 deg R at this high thrust condition. Water and/or oxygen injection are simple and straightforward strategies to improve engine performance and they will add little to engine weight. However, if large Mach number and thrust increases are required, liquid flows become significant, so that operation at these conditions will necessarily be of short duration.
The Electrical Systems Development Branch of the Power Technology Division at the NASA Lewis Research Center in Cleveland, Ohio is designing a Power Management and Distribution (PMAD) System for the Air Force's Integrated Solar Upper Stage (ISUS) Engine Ground Test Demonstration (EGD). The ISUS program uses solar-thermal propulsion to perform orbit transfers from Low Earth Orbit (LEO) to Geosynchronous Orbit (GEO) and from LEO to Molnya. The ISUS uses the same energy conversion receiver to perform the LEO to High Earth Orbit (HEO) transfer and to generate on-orbit electric power for the payloads. On-orbit power generation is accomplished via two solar concentrators heating a dual-cavity graphite-core which has Thermionic Diodes (TMD's) encircling each cavity. The graphite core and concentrators together are called the Receiver and Concentrator (RAC). The TDM-emitters reach peak temperatures of approximately 2200K, and the TID-collectors are run at approximately 1000K. Because of the high Specific Impulse (I(sup sp)) of solar thermal propulsion relative to chemical propulsion, and because a common bus is used for communications, GN&C, power, etc., a substantial increase in payload weight is possible. This potentially allows for a stepdown in the required launch vehicle size or class for similar payload weight using conventional chemical propulsion and a separate spacecraft bus. The ISUS power system is to provide 1000W(sub e) at 28+/-6V(sub dc) to the payload/spacecraft from a maximum TID generation capability of 1070W(sub e) at 2200K. Producing power with this quality, protecting the spacecraft from electrical faults and accommodating operational constraints of the TID's are the responsibilities of the PMAD system. The design strategy and system options examined along with the proposed designs for the Flight and EGD configurations are discussed herein.
The Descent Stage Propulsion System (DSPS) is the most challenging and complex propulsion system ever built at JPL. Performance requirements, such as the entry Reaction Control System (RCS) requirements, and the terminal descent requirements (3300 N maximum thrust and approximately 835,000 N-s total impulse in less than a minute), required a large amount of propellant and a large number of components for a spacecraft that had to fit in a 4.5 meter aeroshell. The size and shape of the aeroshell, along with the envelope of the stowed rover, limited the configuration options for the Descent Stage structure. The configuration and mass constraints of the Descent Stage structure, along with performance requirements, drove the configuration of the DSPS. This paper will examine some of the challenges encountered and solutions developed during the fabrication, assembly, and test of the DSPS.
With the technology currently available for the manufacture of propellants, it is possible to control the variance of the total specific impulse obtained from the rocket boosters to within approximately five percent. Though at first inspection this may appear to be a reasonable amount of control, when it is considered that any uncertainty in the total kinetic energy delivered to the spacecraft translates into a design with less total usable payload, even this degree of uncertainty becomes unacceptable. There is strong motivation to control the variance in the specific impulse of the shuttle's solid boosters. Any small gains in the predictability and reliability of the booster would lead to a very substantial payoff in earth-to-orbit payload. The purpose of this study is to examine one aspect of the manufacture of solid propellants, namely, the mixing process. The traditional approach of computational fluid mechanics is notoriously complex and time consuming. Certain simplifications are made, yet certain fundamental aspects of the mixing process are investigated as a whole. It is possible to consider a mixing process in a mathematical sense as an operator, F, which maps a domain back upon itself. An operator which demonstrates good mixing should be able to spread any subset of the domain completely and evenly throughout the whole domain by successive applications of the mixing operator, F. Two and three dimensional models are developed and graphical visualization two and three dimensional mixing processes are presented.
As a part of a high energy density materials (HEDM) development, the hot fire tests for Quadricyclane, 1,7 Octadiyne, AFRL-1, Biclopropylidene, and CINCH (Dimethyl amino ethyl azide) have been conducted at NASA/MSFC. The first 4 materials for this task are provided from Air Force Research Laboratory at Edward Air Force Base and US Army provided CINCH. The performance of these fuels is compared with RP-1. The preliminary results of these tests are presented. The preliminary results of Quadricyclane tests indicate that the specific impulse and c-star efficiency for quadricyclane at the mixture ratio 1.94 are approximately 5 sec and 105 ft/sec better than the RP-1 at mixture ratio 1.9. The 1,7 Octadiyne test indicate that the specific impulse and c-star efficiency at the mixture ratio 2.1 are approximately -1 sec and 89 ft/sec differ than the RP-1 at mixture ratio 2.04. The Quadricyclane soot buildup at the combustor is a little more than RP-1, but detail study of soot formation is not considered at this time. There was no visual soot buildup for the 1,7 Octadiyne and AFRL-1.
Optimal multi-impulse trajectories were investigated as a nominal about which asymptotic expansion was used to obtain approximations of optimal low thrust trajectories. The work consisted of the analysis and description of an optimal 3-impulse trajectory program. A patched-conic trajectory model was specifically designed for compatibility with the subsequent addition of the low thrust expansion approximation.
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.