Showing papers in "Journal of Spacecraft and Rockets in 1979"
TL;DR: In this article, an experimental apparatus designed for measurement of erosive burning rates at crossflow velocities up to Mach 1 has been used to determine the burning characteristics of seven propellant formulations with systematically varied properties.
Abstract: : An experimental apparatus designed for measurement of erosive burning rates at crossflow velocities up to Mach 1 has been used to determine the erosive burning characteristics of seven propellant formulations with systematically varied properties. A composite propellant erosive burning model based on the bending of columnar diffusion flames gives reasonably good agreement with the measured erosive burning data over a wide range of conditions, breaking down only in regions where the fuel-oxidizer gas stream mixing does not control burning rate. Propellant base (no-crossflow) burning rate is found to have a predominant effect on sensitivity to crossflow (higher- burning-rate formulations being considerably less sensitive) whether the base burning rate differences are produced by oxidizer particle size variation, oxidizer/fuel ratio variation, or use of catalysts. Comparison of erosive burning predictions using the erosive burning model described herein with flow profiles expected to prevail in the test apparatus to predictions using profiles believed to exist in cylindrically-perforated motor grains indicate that erosive burning may be considerably less for a given mainstream crossflow velocity in such a motor than in the typical erosive burning test apparatus, a result quite important to extrapolation of test apparatus erosive burning data to actual motor conditions.
33 citations
TL;DR: In this article, a simplified solid-propellant rocket motor is analyzed as a long hollow elastic cylinder in a thin case (plane strain), and the daily probabilities of failure are determined from a stress-strength interference-based analysis as functions of time and are summed up to produce the probability of failure at the end of the service life.
Abstract: Solid-propellant rocket motors are frequently stored in the field without environmental protection; hence they are subjected to variable thermal stresses and material degradation due to aging and fatigue. Temperature variations are modeled as narrow-band random processes; thermal stresses in the motor will exhibit similar characteristics. Because material properties are statistically distributed, the probability that the thermal stress exceeds the strength of the propellant is synonymous with the probability of failure. A simplified solidpropellant rocket motor is analyzed as a long hollow elastic cylinder in a thin case (plane strain). The daily probabilities of failure are determined from a stress-strength interference-t ype analysis as functions of time and are summed up to produce the probability of failure at the end of the service life. The additional complexities of viscoelasticity, aging, and loads other than thermal will be introduced in subsequent analyses.
32 citations
32 citations
TL;DR: The methodology needed to optimize the design of a scramjet engine using the unified cycle analysis is presented and a specific example applying this optimization procedure to a missile design is also included.
Abstract: Although the concept of using a supersonic combustion ramjet engine to propel a vehicle at hypersonic speeds through the atmosphere has been thoroughly established over the past two decades, and analyses for determining the performance of each of its components documented, no complete engine cycle analysis has been presented to date. The purpose of this paper is to review and discuss the individual component analyses previously developed, combine them into a unified cycle analysis, and present the methodology needed to optimize the design of a scramjet engine using the unified cycle analysis. A specific example applying this optimization procedure to a missile design is also included.
31 citations
TL;DR: In this article, a WIND-TUNNEL test was conducted to determine the support interference on an ogive-cylinder model at high angles of attack in transonic flow, supported by either a base-mounted sting or a strut attached to the leeside of the model.
Abstract: A WIND-TUNNEL test was conducted to determine the support interference on an ogive-cylinder model at high angles of attack in transonic flow. The model was supported by either a base-mounted sting or a strut attached to the leeside of the model. The strut support acted as a splitter plate and generally reduced the normal-force coefficient, while the sting support increased the normal-force coefficient slightly. The support interference diminished with increasing Mach number. A simple algebraic method of estimating support interference was used. Two semiempirical methods for calculation of aerodynamic coefficients were compared with test results.
28 citations
TL;DR: In this article, the authors present a review of techniques and devices that have proved beneficial in meeting aerodynamic simulation requirements for Ramjet engine ground test requirements, including Vitiated air heaters, for direct-connect and freejet tests.
Abstract: Simulation of flow conditions within the operating regime of ramjet engines requires unique ground test capabilities. The blowdown facility is an economical method of meeting the high air mass flow and pressure ratio requirements. Vitiated air heaters provide a flexible and cost-effective method of simulating trajectory temperature variation. Adequate simulation of the inlet flow conditions in freejet tests plays an important role in engine development. Ramjet engine ground test requirements are discussed. Methods of simulating the applicable parameters in direct-connect and freejet tests are reviewed. Techniques and devices that have proved beneficial in meeting aerodynamic simulation requirements are described.
26 citations
TL;DR: A preliminary cost analysis indicates that electric propulsion greatly reduces launch vehicle requirements and would be competitive with chemical propulsion if the technology of power generation systems advances to where reusability can be achieved at low cost.
Abstract: Techniques for sizing electrically or chemically propelled orbit transfer vehicles and analyzing fleet requirements are used in a comparative analysis of the two concepts for various levels of traffic to geosynchronous orbit. The vehicle masses, fuel requirements, and fleet sizes are determined and translated into launch vehicle payload requirements. Technology projections beyond normal growth are made and their effect on the comparative advantages of the concepts is determined. A preliminary cost analysis indicates that electric propulsion greatly reduces launch vehicle requirements and would be competitive with chemical propulsion if the technology of power generation systems advances to where reusability can be achieved at low cost.
23 citations
TL;DR: In this paper, solid hydrazones obtained by the reaction of various aldehydes with phenyl and dimethylhydrazines have been found to be hypergolic with nitric acid and the observed ignition delays have been accounted for by assuming the role of nitration, neutralization and oxidation reactions occurring simultaneously in the preignition stage.
Abstract: Solid hydrazones obtained by the reaction of various aldehydes with phenyl and dimethylhydrazines have been found to be hypergolic with nitric acid. The hypergolicity appears to be related to the chemical structure of these compounds. The observed ignition delays have been accounted for by assuming the role of nitration, neutralization and oxidation reactions occurring simultaneously in the preignition stage. p-Nitrobenzoic acid and benzoic acid have been isolated as preignition reaction intermediates in the benzaldehydephenylhydrazonenitric acid system
22 citations
20 citations
TL;DR: In this article, an axial-groove liquid-trap heat pipe diode with a forward-mode performance of nearly 90 Wm, when ammonia is used as the working fluid at 20°C.
Abstract: An all-aluminum axial-groove liquid-trap heat pipe diode, 470 mm long and 10 mm o.d., has been developed with a forward-mode performance of nearly 90 Wm, when ammonia is used as the working fluid at 20°C. The diode is bendable, of simple design, and of reliable performance. A mathematical model, based on an energy balance for evaporator and trap, has been developed for predicting the transient shutdown of the diode. Theroretical predictions and experimental results are in good agreement. The time for complete shutdown of the diode is on the order of 20 min. The respective shutdown energy is about 4 Wh. A reverse-mode heat flow of about 1.5 W has been measured. Thereby, a turndown ratio of about 300 has been established.
20 citations
TL;DR: The concept of utilizing the space shuttle orbiter as an aerodynamic flight research vehicle is discussed in this paper, where the orbiter's planned flight frequency and its complex flight control system provide an unprecedented flight research potential.
Abstract: The concept of utilizing the space shuttle orbiter as an aerodynamic flight research vehicle is discussed. The orbiter's planned flight frequency and its complex flight control system provide an unprecedented flight research potential. This paper defines the orbiter's flight environment and applicable baseline systems, their capabilities and limitations, as well as those instrument systems required to augment the baseline capability. These required systems, which are being developed under NASA's Orbiter Experiments Program (OEX), are the aerodynamic coefficient indentification package (ACIP), shuttle entry air data system (SEADS), and the shuttle upper atmosphere mass spectrometer (SUMS). Finally, the need for and capability of launching payloads from the orbiter to extend the research potential beyond the orbiter configuration and/or environment is defined.
TL;DR: A magnetoplasmadynamic arcjet is scheduled to be on board the first Spacelab to perform space experiments with particle accelerators as mentioned in this paper, and the results of these tests as well as the hardware design are presented.
Abstract: A magnetoplasmadynamic arcjet is scheduled to be on board the first Spacelab to perform space experiments with particle accelerators. Major scientific missions to be carried out by the arcjet are 1) spacecraft charge neutralization with an electron gun simultaneously operated, 2) plasmadynamics in the background geomagnetic field, and 3) excitation of artificial airglow. Space-chamber tests were carried out twice, in Dec. 1976 and Nov. 1977. The results of these tests as well as the hardware design are presented.
TL;DR: A method wherein the launch vehicle/rigid payload interface accelerations are modified to account for the payload elasticity is developed, which requires the updating of the system normal modes to accounts for payload changes, but does not require a complete transient solution using the composite system model.
Abstract: In the design/analysis process of a payload structural system, the accelerations at the payload/launch vehicle interface obtained from a system analysis using a rigid payload are often used as the input forcing function to the elastic payload to obtain structural design loads. Such an analysis is at best an approximation since the elastic coupling effects are neglected. This paper develops a method wherein the launch vehicle/rigid payload interface accelerations are modified to account for the payload elasticity. The advantage of the proposed method, which is exact to the extent that the physical system can be described by a truncated set of generalized coordinates, is that the complete design/analysis process can be performed within the organization responsible for the payload design. The method requires the updating of the system normal modes to account for payload changes, but does not require a complete transient solution using the composite system model. An application to a real complex structure, the Viking Spacecraft System, is given.
TL;DR: An intercept chart is developed for use in an unaided EVA environment and a brief error analysis is presented along with a series of charts which provide a simple method of estimating the effects of errors in initial conditions on terminal position and velocity.
Abstract: In order to provide a margin of safety for astronauts working untethered in the vicinity of the space shuttle, it is necessary to provide for them a means to determine conditions which would return them to the shuttle via an intercept and rendezvous maneuver An intercept chart is developed for use in an unaided EVA environment Given the in-plane azimuth angle, distance away, and time to intercept, the chart provides the direction and magnitude of A V required for rendezvous The chart is valid for nominally circular orbits of all altitudes Additional charts are presented which provide closing velocity information as well as total AK requirements necessary for planning rendezvous maneuvers Finally, a brief error analysis is presented along with a series of charts which provide a simple method of estimating the effects of errors in initial conditions on terminal position and velocity
TL;DR: In this paper, the effect of interactions of spacecraft-generated plasmas and high voltage solar array components on an advanced solar electric propulsion system was investigated and the means for reducing the parasitic currents collected by the solar arrays were considered.
Abstract: Calculations are made of the effect of interactions of spacecraft-generated plasmas and high voltage solar array components on an advanced Solar Electric Propulsion system. The plasma consists of mercury ions and electrons resulting from the operation of ion thrusters and associated hollow cathode neutralizers. Because large areas of the solar array are at high potential and not completely insulated from the surrounding plasma, the array can, under some conditions, collect excessive electron currents. Results are given for the parasitic currents collected by the solar arrays and means for reducing these currents are considered.
TL;DR: Romodanova, L. D. and Pokhil, P. K., this article, "Action of Silica on the Burning Rates of Ammonium Perchlorate Compositions," Fizika Goreniya, I. Vzryva and I. Cohen Nir, E., ''Combustion of Powered Metals in Contact with a Solid Oxidiser," Thirteenth Symposium (International) on Combustion, The Combustions Institute, Pittsburgh, Pa., 1971, pp. 1019-1029.
Abstract: References 1 Romodanova, L. D. and Pokhil, P. K., "Action of Silica on the Burning Rates of Ammonium Perchlorate Compositions," Fizika Goreniya, I. Vzryva., Vol. 6, No. 3, 1970, pp. 285-288. Cohen Nir, E., ''Combustion of Powered Metals in Contact with a Solid Oxidiser," Thirteenth Symposium (International) on Combustion, The Combustion Institute, Pittsburgh, Pa., 1971, pp. 1019-1029. Boggs, T. L. and Zurn, D. E., "Ammonium Perchlorate Combustion: Effects of Sample Preparation; Ingredient Type; and Pressure, Temperature and Acceleration Environments," Combustion Science and Technology, Vol. 7, 1973, pp. 177-183. Rossini, F. D., Wagman, D. D., Evans, W. H., Levine, S., and Jaffe, I., "Selected Values of Chemical Thermodynamic Properties," National Bureau of Standards, 1952.
TL;DR: In this paper, a laboratory-scale test program was conducted to investigate candidate mechanisms that could explain the unacceptably low bubble pressures that were obtained during the reaction control system (RCS) tank tests.
Abstract: A laboratory-scale test program was conducted to investigate candidate mechanisms that could explain the unacceptably low bubble pressures that were obtained during the reaction control system (RCS) tank tests. It was concluded that liquid evaporation from the fine-mesh screens was the cause of the observed test results. Correlations of the data were developed for two mechanisms of liquid evaporation from the screens: evaporation caused by temperature gradients and evaporation caused by unsaturated pressurant gas. These correlations identify the environmental requirements for conducting successful RCS tank bubble pressure tests.
TL;DR: In this paper, the authors describe the design of these experiments, the instrumentation, experimental results, and related observations during array deployment and ground tests, and compare with a theoretical model.
Abstract: Hermes is a geostationary communications satellite equipped with large flexible solar arrays instrumented for experimental purposes. A program for obtaining flight data on structural dynamics was part of the mission, including special in-flight experiments. This paper describes the design of these experiments, the instrumentation, experimental results, and related observations during array deployment and ground tests. Flight results were obtained on natural frequencies, damping factors, and steady-state resonant amplitudes. Finally, comparisons with a theoretical model are given. Generally, excellent agreement was found between measurments and predicted values. However, damping factors measured were higher than expected from ground tests.
TL;DR: The economic effect of incorporating dualfuel propulsion in advanced vehicles is shown by comparing several dual-fuel propulsion systems compared to a baseline hydrogen and oxygen system.
Abstract: Econometric analyses of advanced Earth-to-orbit vehicles indicate that there are economic benefits from the development of new vehicles beyond the space shuttle as traffic increases. Vehicle studies indicate the advantage of the dual-fuel propulsion in single-stage vehicles. This paper shows the economic effect of incorporating dualfuel propulsion in advanced vehicles. Several dual-fuel propulsion systems are compared to a baseline hydrogen and oxygen system.
TL;DR: The original 12-cm hexagonal magneto-electrostatic containment discharge chamber described by Moore in 1969 has been optimized for argon and xenon operation in this paper, achieving a mass utilization of 67-80% at keeper-plus-main discharge energy consumptions of 200-458 eV/ion, respectively.
Abstract: The original 12-cm hexagonal magneto-electrostatic containment discharge chamber described by Moore in 1969 has been optimized for argon and xenon operation. Argon mass utilization efficiencies of 67-80% were achieved at keeper-plus-main discharge energy consumptions of 200-458 eV/ion, respectively. Xenon performance of 85-96% mass utilization was realized at 200-350 eV/ion. The paper discusses the optimization process and test results.
TL;DR: In this paper, the internal ballistic effects of combined radial and circumferential grain temperature gradients are evaluated theoretically for the Space Shuttle solid rocket motors (SRM's), and a simplified approach is devised for representing with closed-form mathematical expressions the temperature distribution resulting from the anticipated thermal history prior to launch.
Abstract: The internal ballistic effects of combined radial and circumferential grain temperature gradients are evaluated theoretically for the Space Shuttle solid rocket motors (SRM's). A simplified approach is devised for representing with closed-form mathematical expressions the temperature distribution resulting from the anticipated thermal history prior to launch. The internal ballistic effects of the gradients are established by use of a mathematical model which permits the propellant burning rate to vary circumferentially. Comparative results are presented for uniform and axisymmetric temperature distributions and the anticipated gradients based on an earlier two-dimensional analysis of the center SRM segment. The thrust imbalance potential of the booster stage is assessed based on the difference in the thermal loading of the individual SRM's of the motor pair which may be encountered in both summer and winter environments at the launch site. Results indicate that grain temperature gradients could cause the thrust imbalance to be approximately 10% higher in the Space Shuttle than the imbalance caused by SRM manufacturing and propellant physical property variability alone.
TL;DR: In this article, the effects of igniter strength, propellant type, deterrent concentration, and projectile motion on the overall transient combustion processes in granular propellant beds were studied experimentally.
Abstract: Catastrophic failure of propulsion and gun systems has been attributed to abnormal combustion within granular propellant charges. The effects of igniter strength, propellant type, deterrent concentration, and projectile motion on the overall transient combustion processes in granular propellant beds were studied experimentally. The results show that igniter strength significantly affects the duration of the induction period and the accelerative behavior of the pressure front traveling through the bed; a weaker igniter causes a more pronounced pressure front acceleration. A large igniter volume was found to reduce the rate of flame-spreading and pressurization processes. Combustion of slightly deterred propellants produced very rapid flame spreading, higher peak pressures, and higher pressurization rates than regularly deterred propellants. Propellant particle geometry was found to greatly affect the rate of total mass consumption within a propellant bed and thereby influence the peak pressures and pressure wave phenomena within the bed.
TL;DR: In this article, a systematic and generalized analysis of thermal control characteristics of movable louvers with a finite number of blades is presented, based on the assumption that the temperature of each blade is uniform and all blades are infinite in length.
Abstract: A systematic and generalized analysis has been made for evaluation of thermal control characteristics of movable louvers with a finite number of blades. The analysis is based on the assumption that the temperature of each blade is uniform and all blades are infinite in length. The modified radiosity approach is used to describe multiple reflections in a louver channel. The formulation is extended to include the radiant interchange due to one, two, or three intervening specular reflections. The image method is employed to estimate generalized view factors expressing fractions of radiant interchange due to specular reflections. A semigray spectral subdivision of surface properties is employed to account for differences in property values for solar and surface radiation. In each spectral interval a reflectance model with both diffuse and specular components is used to approximate the nondiffuse character of surface reflectance. An algorithm based on three-dimensional analytical geometry is developed for estimating the intensity of the solar field in a louver channel. The analysis yields closed-form expressions with which one can calculate effective emittance, solar rejection capability, and base surface temperature required for active thermal control. Numerical results of parametric studies are presented to aid the optimal design of movable louvers for space use. The results on effective absorptance and blade temperature are compared with experimental results obtained for thermal control louvers used on the ATS-F&G spacecraft. Agreement is good in spite of some differences between model and reality.
TL;DR: In this paper, the authors used the linear cumulative damage theory and the concept of failure area to derive the relationship between the stress intensity factor and the crack propagation rate in polyurethane and carboxyterminated polybutadiene propellants.
Abstract: Determination of the relationships that govern crack propagation in propellant material is essential for predicting the failure of grains. Until now, the most useful crack propagation law was the Paris one which gives a power relation between the stress intensity factor and the crack propagation rate. However, this relation appears to be inadequate for a correct representation of crack propagation in viscoelastic materials; by using the linear cumulative damage theory and the concept of failure area, a modified law may be chosen: aT da/dt = A/tm. This proposed law allows a further reduction of the dispersion of the results for tensile experiments on uniaxial precracked specimens. The propellants used in this study are polyurethane and carboxyterminated polybutadiene propellants. Nomenclature A,A' = constant (crack propagation) a = crack length aT =W.L.F. shift factor b =half of the width of the specimen D = damage da/dt — crack velocity El = constant Erd = relaxation modulus k = constant (damage law) Kj = stress-intensity factor /V = number of cycles in a fatigue experiment n = constant (power law) p = constant (damage law) q — constant (crack propagation) R = strain rate T = temperature t = time ta — characteristic time tm =time corresponding to the maximum stress om during a uniaxial tensile experiment a. — length of failure zone \l/ = geometric factor a = applied stress o0 = constant (power law) om = maximum stress during a uniaxial tensile experiment F = fracture energy
TL;DR: Fluctuating pressure measurements at M = 5.0 are presented for turned (axisymmetric) nose shapes that represent approximations of cross sections (at various meridians) of the recovered graphite nosetip of the NRV.
Abstract: Fluctuating pressure measurements at M = 5.0 are presented for turned (axisymmetric) nose shapes that represent approximations of cross sections (at various meridians) of the recovered graphite nosetip of the Nosetip Recovery Vehicle (NRV). Selected fluctuating pressure measurements are also presented for a replica of the NRV nosetip and a replica of a wind tunnel ablated camphor nosetip that roughly resembles the NRV nosetip. The greatest fluctuating pressures occurred under imbedded shock waves and/or the reattachment shock waves of imbedded regions of separated flow. The non-dimensionalized power spectra under the reattachment shocks are in good agreement with similar measurements for spiked cones and a drag-reduction spike. Evidence of a Reynolds number sensitive resonant peak was observed in the spectra under reattachment on the NRV nose due to surface irregularities.
TL;DR: In this paper, a concept for a lightweight space-based orbital transfer vehicle (OTV) featuring thin, spherical, pressuredesigned aluminum liquid-oxygen and liquid-hydrogen tanks and a truss structure of composite materials is used as a baseline design for a large-cargo OTV.
Abstract: A concept for a lightweight space-based orbital transfer vehicle (OTV) featuring thin, spherical, pressuredesigned aluminum liquid-oxygen and liquid-hydrogen tanks and a truss structure of composite materials is used as a baseline design for a large-cargo OTV. Vehicle sizing, fleet analysis, and parametric cost analysis are used to evaluate the effects of engine technology, vehicle staging, and high-thrust vs. low-thrust transfer. Results indicate that Earth-to-orbit launch costs and OTV engine performance are strong drivers in orbital transportation cost and that there is no significant benefit in staging vehicles. At the low values of Earth-to-orbi t cost representative of advanced launch vehicles, low-thrust and high-thrust OTV's are competitive.