Showing papers in "Journal of Spacecraft and Rockets in 1995"
TL;DR: In this paper, the authors analyzed the energy fluences measured at GOES-7 and METEOSAT-3 and calculated the critical directed energy fluence, which is consistent with the theoretical breakdown threshold and CRRES data characterizing the discharge process; it fits with an effective shielding thickness of less than 0.2 mm.
Abstract: When the ANIK El and £2 communications spacecraft suffered serious failures of their momentum wheel control systems, it was postulated that the satellites were subjected to bulk (internal dielectric) charging followed by discharge that disabled key circuitry. This paper effectively confirms the hypothesis by linking the events to a well-established pattern of operational anomalies on another spacecraft. Since March 1991, a commercial geosynchronous satellite has experienced over 50 specific, but relatively trivial, mode switches. These are analyzed in relation to energetic electron fluences measured at GOES-7 and METEOSAT-3; without exception they coincide with periods of relatively high flux. Combining the measurements, the critical directed energy fluence is calculated to be 1 x 10 MeV cm". This figure is consistent with the theoretical breakdown threshold and CRRES data characterizing the discharge process; it fits with an effective shielding thickness of less than 0.2 mm of aluminum. Three switches in the days preceding the ANIK failures strongly argue for a common explanation, although it may be impossible to even identify actual discharge sites. Modern communications spacecraft employ tried and tested techniques for electrostatic discharge protection, but it is clear that the recognized hazard of internal dielectric charging has often been underestimated and shielding guidelines overlooked.
157 citations
TL;DR: In this paper, an application of approximate heat balance integral (HBI) methods to more realistic models of material response, specifically, for carbon-carbon and carbon-phenolic heat shields on reentry vehicles, is presented.
Abstract: To predict ablation, charring, and erosion of heat-shield materials, approximate heat balance integral (HBI) methods offer speed and versatility; however, traditional HBI articles treat only simple, idealized models of material response. This paper reviews application of HBI methods to more realistic models of material response, specifically, for carbon-carbon and carbon-phenolic heat shields on reentry vehicles. The review shows that HBI successfully extends to most such simulations of ablation, charring, and erosion in hypersonic flow, but unexpected problems can crop up and trade-offs exist. Pertinent material models are also summarized, including efficient expressions that fit material thermal properties and carbon-air thermochemical ablation functions.
113 citations
TL;DR: In this paper, the authors investigated the vibrational disturbances of the Hubble Space Telescope that were discovered soon after deployment in orbit, and found that the disturbances were particularly evident during orbital day-night crossings, and that the magnitude of the disturbances was considerably larger than the design jitter requirement.
Abstract: The investigation of the vibrational disturbances of the Hubble Space Telescope that were discovered soon after deployment in orbit is described in detail. It was found that the disturbances were particularly evident during orbital day-night crossings, and that the magnitude of the disturbances was considerably larger than the design jitter requirement. This paper describes the process by which the vibrations were characterized and isolated to a particular mechanism. The analysis of the flight data and comparisons with computer simulation results showed that the source of the disturbances was the thermally driven deformation of the solar arrays in conjunction with frictional effects in the array mechanisms. The control system was successfully modified to attenuate the disturbances to tolerable levels pending mechanical and thermal redesign of the solar arrays. The new arrays were installed during the first Space Telescope servicing mission, and in combination with the enhanced control system algorithm reduced the disturbances to satisfactory levels.
112 citations
TL;DR: In this article, a six-degree-of-freedom trajectory analysis and aerodynamic characteristic assessment are performed to demonstrate vehicle flyability and quantify the effect that each of numerous uncertainties has upon the nominal mission profile.
Abstract: The Mars Pathfinder mission provides the next opportunity for scientific exploration of the surface of Mars following a 7.6 km/s direct entry. In support of this effort, a six-degree-of-freedom trajectory analysis and aerodynamic characteristic assessment are performed to demonstrate vehicle flyability and to quantify the effect that each of numerous uncertainties has upon the nominal mission profile. The entry vehicle is shown to be aerodynamically stable over a large portion of its atmospheric flight. Two low angle-of-attack static instabilities (freestream velocities of about 6.5 and 3.5 km/s) and a low angle-of-attack dynamic instability (supersonic) are identified and shown to cause bounded increases in vehicle attitude. The effects of center-of-gravity placement, entry attitude, vehicle roll rate, aerodynamic misprediction, and atmospheric uncertainty on the vehicle attitude profile and parachute deployment conditions are quantified. A Monte Carlo analysis is performed to statistically assess the combined impact of multiple off-nominal conditions on the nominal flight characteristics. These results suggest that there is a 99.7% probability that the peak attitude throughout the entry will be less than 8.5 deg, the peak heating attitude will be below 6.2 deg, and the attitude at parachute deployment will be less than 3.9 deg.
101 citations
TL;DR: In this article, the effects of the spike length, Mach number, and angle of attack on the supersonic flow were examined using three-dimensional thin-layer compressible Navier-Stokes equations.
Abstract: In supersonic flow, a spike attached to the nose reduces the drag of a blunt body. In this paper, supersonic flows around a spiked blunt body are numerically simulated to examine the effects of the spike length, Mach number, and angle of attack. Three-dimensional thin-layer compressible Navier-Stokes equations are solved using a highresolution upwind scheme with LU-ADI time-integration algorithm. The computed results show that the drag of the spiked blunt body is significantly influenced by the spike length, Mach number, and angle of attack. Scales of the separated region are not significantly influenced by the freestream Mach number. For the spiked blunt body at angle of attack, the flowfield becomes complex with spiral flows. The computed results are in reasonable agreement with experimental data.
101 citations
TL;DR: In this paper, two multidisciplinary optimization strategies in which the analyses are integrated in different manners are implemented and evaluated for solution of a single-stage-to-orbit launch-vehicle design problem.
Abstract: The investigation focuses on development of a rapid multidisciplinary analysis and optimization capability for launch-vehicle design. Two multidisciplinary optimization strategies in which the analyses are integrated in different manners are implemented and evaluated for solution of a single-stage-to-orbit launch-vehicle design problem. Weights and sizing, propulsion, and trajectory issues are directly addressed in each optimization process. Additionally, the need to maintain a consistent vehicle model across the disciplines is discussed. Both solution strategies were shown to obtain similar solutions from two different starting points. These solutions suggests that a dual-fuel, single-stage-to-orbit vehicle with a dry weight of approximately 1.927 x 10(exp 5)lb, gross liftoff weight of 2.165 x 10(exp 6)lb, and length of 181 ft is attainable. A comparison of the two approaches demonstrates that treatment or disciplinary coupling has a direct effect on optimization convergence and the required computational effort. In comparison with the first solution strategy, which is of the general form typically used within the launch vehicle design community at present, the second optimization approach is shown to he 3-4 times more computationally efficient.
90 citations
TL;DR: In this article, the aerothermodynamics of the Mars Pathfinder vehicle entering the Mars atmosphere at the maximum heating and maximum deceleration points in its trajectory are described. Ablating and nonablating boundary conditions are developed which produce maximum recombination of CO2 on the surface.
Abstract: A computational approach is used to describe the aerothermodynamics of the Mars Pathfinder vehicle entering the Mars atmosphere at the maximum heating and maximum deceleration points in its trajectory. Ablating and nonablating boundary conditions are developed which produce maximum recombination of CO2 on the surface. For the maximum heating trajectory point, an axisymmetric, nonablating calculation predicts a stagnation-point value for the convective heating of 115 W/cm(exp 2). Radiative heating estimates predict an additional 5-12 W/cm(exp 2) at the stagnation point. Peak convective heating on the afterbody occurs on the vehicle's flat stern with a value of 5.9% of the stagnation value. The forebody flow exhibits chemical nonequilibrium behavior, and the flow is frozen in the near wake. Including ablation injection on the forebody lowers the stagnation-point convective heating 18%.
82 citations
TL;DR: A comprehensive review of experimental base pressure and base heating data related to supersonic and hypersonic flight vehicles is presented in this article, where a series of internally consistent, empirical predictions are developed for planar and axisymmetric geometries (wedges, cones, and cylinders).
Abstract: A comprehensive review of experimental base pressure and base heating data related to supersonic and hypersonic flight vehicles is presented. Particular attention is paid to free-flight data as well as wind-tunnel data for models without rear sting support. Using theoretically based correlation parameters, a series of internally consistent, empirical predictions are developed for planar and axisymmetric geometries (wedges, cones, and cylinders). These equations encompass the speed range from low supersonic to hypersonic flow and laminar and turbulent forebody boundary layers. A wide range of cone and wedge angles and cone bluntness ratios is included in the data base used to develop the correlations. The present investigation also includes an analysis of the effect of the angle of attack and the specific-heat ratio of the gas. Angle-of-attack effects are considered on sharp and blunted cones and cylindrical afterbodies.
76 citations
TL;DR: In this article, the authors demonstrate that the computational fluid-dynamics code interfacing with the material response code can be directly applied to the design of thermal protection systems of spacecraft.
Abstract: The Mars Pathfinder probe will enter the Martian atmosphere at a relative velocity of 7.65 km/s. The 2.65-m-diam vehicle has a blunted, 70-deg-half-angle, conical forebody aerobrake. Axisymmetric time-dependent calculations have been made using Gauss-Seidel implicit aerothermodynamic Navier-Stokes code with thermochemical surface conditions and a program to calculate the charring-material thermal response and ablation for heating analysis and heat-shield material sizing. The two codes are loosely coupled. The flovvfield and convective heat-transfer coefficients are computed using the flowfield code with species balance conditions for an ablating surface. The timedependent in-depth conduction with surface blowing is simulated using the material response code with complete surface energy-balance conditions. This is the first study demonstrating that the computational fluid-dynamics code interfacing with the material response code can be directly applied to the design of thermal protection systems of spacecraft. The heat-shield material is SLA-561V. The solutions, including the flowfield, surface heat fluxes and temperature distributions, pyrolysis-gas blowing rates, in-depth temperature history, and minimum heat-shield thicknesses over the aeroshell forebody, are presented and discussed in detail. The predicted heat-shield mass is about 20 kg.
72 citations
TL;DR: The evaluation of oxygen interactions with materials III space-flight experiment was developed to obtain benchmark atomic-oxygen reactivity data and was conducted during Space Transportation System Mission 46 as mentioned in this paper, where mass-spectrometric measurements of gaseous products formed by O-atom reaction with 13C-labeled Kapton™ revealed CO, CO2, H2O, NO, and NCh.
Abstract: The Evaluation of Oxygen Interactions with Materials III space-flight experiment was developed to obtain benchmark atomic-oxygen reactivity data and was conducted during Space Transportation System Mission 46. We present an overview of the flight experiment and the results of the Lyndon B. Johnson Space Center polymer chemistry and mass-spectrometer-carousel experiments. Mass-spectrometric measurements of gaseous products formed by O-atom reaction with 13C-labeled Kapton™ revealed CO, CO2, H2O, NO, and NCh. By operating the mass spectrometer to detect naturally occurring ionospheric species, we characterized the ambient ionosphere at various times during the flight experiment and detected the gaseous reaction products formed when ambient ions interacted with the 13C Kapton carousel sector. Direct comparison of the results of on-orbit O-atom exposures with those conducted in ground-based laboratory systems, which provide known O-atom fluences and translational energies, demonstrated the strong translational-energy dependence of O-atom reactions with a variety of polymers. A line-of-centers reactive scattering model was shown to provide a reasonably accurate description of the translational-energy dependence of polymer reactions with O atoms at high atom kinetic energies, and a Beckerle-Ceyer model provided an accurate description of O-atom reactivity over a three-order-of -magnitude range in translational energy and a four-order-of-magnitude range in reaction efficiency. Postflight studies of the polymer samples by x-ray photoelectron spectroscopy and infrared spectroscopy demonstrate that O-atom attack is confined to the near-surface region of the sample, that is, within 50 to 100 A of the surface.
71 citations
TL;DR: The evaluation of oxygen interaction with materials (EOIM-IH) flight experiment was developed to obtain benchmark atomic-oxygen-material reactivity data The experiment was conducted during Space Shuttle mission 46, July 31 to August 7, 1992 Quantitative interpretation of the materials reactivity measurements requires a complete and accurate definition of the space environment exposure, including the thermal history of the payload, the solar ultraviolet exposure, the atomicoxygen fluence and any spacecraft outgassing and contamination effects as mentioned in this paper.
Abstract: The Evaluation of Oxygen Interactions with Materials III (EOIM-IH) flight experiment was developed to obtain benchmark atomic-oxygen-material reactivity data The experiment was conducted during Space Shuttle mission 46, July 31 to August 7, 1992 Quantitative interpretation of the materials reactivity measurements requires a complete and accurate definition of the space environment exposure, including the thermal history of the payload, the solar ultraviolet exposure, the atomic-oxygen fluence, and any spacecraft outgassing and contamination effects The thermal history of the payload was measured using 11 thermocouple sensors placed behind selected samples and on the EOIM-III payload structure The solar ultraviolet exposure history of the EOIM-IH payload was determined by analysis of the as-flown orbit and vehicle attitude combined with daily average solar ultraviolet and vacuum ultraviolet fluxes The atomic-oxygen fluence was assessed in three ways First, the O-atom fluence was calculated using a program that incorporates the MSIS-86 atmospheric model, the as-flown Space Shuttle trajectory, and solar activity parameters Second, it was estimated directly from Kapton film erosion Third, ambient O-atom measurements were made using the quadrupole mass spectrometer on the EOIM-III payload As of this writing, our best estimate of the O-atom fluence is (23 ± 03) X1020 atoms/cm2 Finally, results of postflight surface analysis of selected samples by x-ray photoelectron spectroscopy indicate low levels of molecular contamination on the payload surface
TL;DR: In this article, a comprehensive analysis of erosion data for materials exposed to low-Earth-orbit space environment, to fast atomic-oxygen beams, and in plasma facilities is shown that different correlations can be found between the thermal and hyperthermal atomicoxygen erosion yield of hydrocarbon polymers and their chemical structure and content.
Abstract: From a comprehensive analysis of erosion data for materials exposed to low-Earth-orbit space environment, to fast atomic-oxygen beams, and in plasma facilities it is shown that different correlations can be found between the thermal and hyperthermal atomic-oxygen erosion yield of hydrocarbon polymers and their chemical structure and content. Correlations have been found of the hyperthermal atomic-oxygen erosion yield of many polymeric materials in flight experiments with their inverse mass density of effective (not bonded with oxygen) carbon atoms, and with their degree of aromaticity. These correlations were used to identify the rate-limiting factors of erosion processes and to predict the erosion rate for polymer-based materials in low Earth orbit. The first correlation was not found in the interaction of thermal atomic oxygen with a number of polymers, and the second is strongly pronounced. The results are explained on the basis of physical and chemical processes affecting differently the erosion rates of polymers by thermal and by fast atomic oxygen. Subthreshold bombardment-induced and -enhanced degradation and surface chemical etching are proposed to be the major mechanisms of erosion by fast atomic oxygen.
TL;DR: This is the first comprehensive review of large spacecraft appendages that are both deployable and retractable to guide the design, development, and ground testing of retractable appendages for future spacecraft.
Abstract: This is the first comprehensive review of large spacecraft appendages that are both deployable and retractable. Its aim is to gather information on retraction-specific issues, to guide the design, development, and ground testing of retractable appendages for future spacecraft. Following a survey of existing retractable booms and masts, solar arrays, and antennas, including examples of special latching systems, an extensive investigation of the in-orbit performance of appendages that were deployed and retracted in space during the last 25 years has been carried out. Remarkably, almost all appendages that had deployed successfully could also be retracted and, despite widespread concern about total reliance on electric motors for deployment and retraction, the evidence shows that practically all electric motors have performed well.
Abstract: A two-temperature model is developed for the description of thermal and chemical nonequilibrium viscous hypersonic flows including ionization. A preferential dissociation model and nonpreferential removal of vibrational and electronic energy are assumed. For weakly ionized flows, an ambipolar diffusion coefficient is introduced to describe ion diffusion. The numerical technique relies on a finite-volume approach based on a second-order accurate Total-Variation-Diminishing formulation that allows for thermal and chemical nonequilibrium effects as well as for ionization. The model has been applied to compute ionizing hypersonic flows over a wedge and a RAM-C geometry. Applications have shown that, for weakly ionized flows, ionization is essentially decoupled from the other field properties. Moreover, the computations show the importance of considering kinetic and diffusive mechanisms fully coupled in order to properly understand the flow features.
TL;DR: In this article, the authors describe the simulator, simulation results, and the Hitchhiker experiment in the context of the Gravity and Magnetic Earth Surveyor subsatellite aerostabilization requirements.
Abstract: The Passive Aerodynamically Stabilized Magnetically Damped Satellite experiment will characterize and demonstrate passive aerodynamic stabilization and passive magnetic hysteresis damping of attitude rates. It is currently scheduled to be deployed on a Shuttle Hitchhiker flight. Although theoretically feasible, aerodynamically induced passive attitude stability represents a technology that has never been substantiated through actual flight experience. The two-week experiment will serve to validate overall performance predictions by the high-fidelity free-molecularflow simulation code developed at the Langley Research Center of NASA. The code can simulate with high fidelity the flight characteristics of a satellite in low Earth orbit. Aerostabilization, if proved, is highly desirable for future missions such as the Gravity and Magnetic Earth Surveyor. This paper describes the simulator, simulation results, and the Hitchhiker experiment in the context of the Gravity and Magnetic Earth Surveyor subsatellite aerostabilization requirements.
TL;DR: In this paper, a response-surface methodology for multidisciplinary design was utilized to optimize the dual-fuel vehicles with respect to several important propulsion-system and vehicle design parameters, in order to achieve minimum empty weight.
Abstract: As part of the United States Advanced Manned Launch System study to determine a follow-on, or complement, to the Space Shuttle, a reusable single-stage-to-orbit concept utilizing dual-fuel rocket propulsion has been examined Several dual-fuel propulsion concepts were investigated These include: a separate-engine concept combining Russian RD-170 kerosene-fueled engines with space shuttle main engine-derivative engines: the kerosene- and hydrogen-fueled Russian RD-701 engine; and a dual-fuel, dual-expander engine Analysis to determine vehicle weight and size characteristics was performed using conceptual-level design techniques A response-surface methodology for multidisciplinary design was utilized to optimize the dual-fuel vehicles with respect to several important propulsion-system and vehicle design parameters, in order to achieve minimum empty weight The tools and methods employed in the analysis process are also summarized In comparison with a reference hydrogen- fueled single-stage vehicle, results showed that the dual-fuel vehicles were from 10 to 30% lower in empty weight for the same payload capability, with the dual-expander engine types showing the greatest potential
TL;DR: In this paper, the acceleration input is notched at the resonance frequencies of the test item in order to limit the shaker force to that predicted for flight since few flight vibratory force data are available, the force limits must be calculated from measurements or analyses of the flight mounting structure and test item mechanical impedances.
Abstract: An improved vibration testing technique has been recently developed and applied to eliminate overtesting caused by the infinite mechanical impedance of the shaker in conventional vibration tests With the new technique, the acceleration input is notched at the resonance frequencies of the test item in order to limit the shaker force to that predicted for flight Since few flight vibratory force data are available, the force limits must be calculated from measurements or analyses of the flight mounting structure and test-item mechanical impedances An improved method of calculating these force limits is derived from evaluating the test-item apparent mass at the coupledsystem resonance frequencies Application of the method to a simple and to a complex coupled oscillator system yields nondimensional results which may be used to calculate limits for future force-limited vibration tests For example, using the simple system results with Q = 50 and equal impedances of the flight mounting structure and test item, the input acceleration amplitude will be notched by a factor of 3125 relative to a conventional test
TL;DR: In this paper, various concepts for advanced manned launch systems (AMLS) are examined for delivery missions to the Space Station and polar orbit, including single and two-stage winged systems with rocket and/or airbreathing propulsion systems.
Abstract: Various concepts for advanced manned launch systems (AMLS) are examined for delivery missions to Space Station and polar orbit. Included are single- and two-stage winged systems with rocket and/or airbreathing propulsion systems. For near-term technologies, two-stage, reusable rocket systems are favored over single-stage rocket or two-stage airbreathing/rocket systems. Advanced technologies enable viable single-stage-to-orbit (SSTO) concepts. Although two-stage rocket systems continue to be lighter in dry weight than SSTOs, advantages in simpler operations may make SSTOs more cost effective over the life cycle. Generally, rocket systems maintain a dry weight advantage over airbreathing systems at the advanced technology levels, but to a lesser degree than when near-term technologies are used. More detailed understanding of vehicle systems and associated ground and flight operations requirements and procedures is essential in determining quantitative discrimination between these latter concepts.
TL;DR: In this paper, a study of the microstructural damage and fracture processes associated with the fracturetoughness behavior of a polymer-matrix composite solid rocket-propellant material was made.
Abstract: A study has been made of the microstructural damage and fracture processes associated with the fracturetoughness behavior of a polymer-matrix composite solid rocket-propellant material. Specifically, nonlinear-elastic fracture-mechanics tests were performed, as a function of displacement rate and temperature (—54 to 71°C), on center-cracked sheet test samples to determine fracture toughness in the form of /-integral resistance curves and viscoelastic resistance curves for the inert propellant H-24; in addition, in situ video imaging was employed to characterize the deformation and interaction between microstructural features and the crack-path morphology. It was found that at the lowest temperatures the increased polymer matrix strength resulted in enhanced cavitation and particle delamination, leading to larger crack-tip fracture process zones and hence to higher fracture toughness. At the high temperatures, the weaker polymer matrix was seen to tear readily and to advance the crack before significant particle delamination or cavitation would occur; this mechanism resulted in small fracture process zones and accounts for the decrease in fracture toughness with increasing temperature. The relationships between the toughness and fracture process in the material are discussed in terms of characterizing parameters for microstructural damage.
TL;DR: In this article, the influence of fuel-grain design on the overall performance of hybrid rocket boosters is investigated, based on steady, one-dimensional compressible flow, including the capability to handle arbitrary wagon-wheel fuel-section designs.
Abstract: A ballistics model has been developed for investigating the influence of fuel-grain design on the overall performance of hybrid rocket boosters. The model, based on steady, one-dimensional compressible flow, includes the capability to handle arbitrary wagon-wheel fuel-section designs. To properly evaluate tradeoffs associated with a booster, stagnation-pressure losses are evaluated and a throttling capability is included. Results of the ballistics calculations are presented for both liquid oxygen and 90% hydrogen peroxide oxidizers assuming hydroxy-terminated polybutadiene as fuel. Vehicle interactions are considered by investigating a design capable of accomplishing the Titan 34D booster mission. Results indicate that the liquid-oxygen systems tend to optimize to a higher number of ports and mass-flux levels than the hydrogen peroxide systems.
TL;DR: The HRDI is a triple-etalon Fabry-Perot interferometer with a resolution of 0.05 cm'1 and very good white-light rejection as discussed by the authors.
Abstract: 3£J) band. The HRDI is a triple-etalon Fabry-Perot interferometer with a resolution of ~0.05 cm'1 and very good white-light rejection. Careful design and calibration has limited systematic errors in the wind determination to less than 5 m/s.
TL;DR: In this paper, the effects of transpiration cooling on boundary-layer growth, heat transfer, and nozzle performance were investigated using a two-dimensional Mach 2.0 contoured nozzle with one contoured wall constructed of sintered stainless steel (2.0-m pore size).
Abstract: The effects of transpiration cooling on boundary-layer growth, heat transfer, and nozzle performance were investigated. A two-dimensional Mach 2.0 contoured nozzle (Re/m = 5.2 x 10 7), with one contoured wall constructed of sintered stainless steel (2.0-^m pore size) was tested. Blowing ratios up to 0.51% of the freestream mass flow were tested using high-frequency-response heat flux and pressure instrumentation. Measurements included wall heat flux and static pressure, as well as exit pitot pressure profiles. Shadowgraph photography was used for flow visualization. A reduction in heat transfer of up to 14% was measured for the highest blowing ratio. On the other hand, large increases in the nozzle exit boundary-layer thickness were found. However, blowing had a minimal effect on the nozzle thrust coefficient and specific impulse. In general, this study highlighted the potential beneficial and adverse effects of transpiration cooling on rocket nozzle cooling.
TL;DR: In this paper, the authors examined the effects of Mach number, Reynolds number, and ratio of specific heat ratio gamma on the nose-up pitching moment of the first entry of the Shuttle Orbiter.
Abstract: During the high-Mach-number, high-altitude portion of the first entry of the Shuttle Orbiter, the vehicle exhibited a nose-up pitching moment relative to preflight prediction of approximately Delta Cm = 0.03. This trim anomaly has been postulated to be due to compressibility, viscous, and/or real-gas (lowered specific heat ratio gamma) effects on basic body pitching moment, body-flap effectiveness, or both. In order to assess the relative contribution of each of these effects, an experimental study was undertaken to examine the effects of Mach number, Reynolds number, and ratio of specific heats. Complementary computational solutions were obtained for wind-tunnel and flight conditions. The primary cause of the anomaly was determined to be lower pressures on the aft windward surface of the Orbiter than deduced from hypersonic wind-tunnel tests with ideal- or near-ideal-gas test flow. The lower pressure levels are a result of the lowering of the flowfield gamma due to high-temperature effects. This phenomenon was accurately simulated in a hypersonic wind tunnel using a heavy gas, which provided a lower, gamma, and was correctly predicted by Navier-Stokes computations using nonequilibrium chemistry.
TL;DR: A new method for the design of inviscid waverider configurations with more general shock geometries than previously possible is presented, utilizing a cross-stream marching scheme for solving the Euler equations.
Abstract: A new method for the design of inviscid waverider configurations with more general shock geometries than previously possible is presented. An arbitrary three-dimensional shock shape is specified as input, and a new inverse method, utilizing a cross-stream marching scheme for solving the Euler equations, is used to generate the postshock flowfield. Unlike most previous studies, this approach allows for the use of nonaxisymmetric shock topologies with nonconstant shock strengths. The problem's ill-posedness is suppressed by reformulating the problem in the proper curvilinear coordinate system. The inverse marching approach is briefly summarized, and details of the waverider design procedure are given. Comparisons of individual waveriders and their flowfields generated by the new algorithm in a few seconds on a workstation with flowfields computed by a direct Euler solver, requiring on the order of an hour of Cray 2 CPU time, demonstrate the accuracy and efficiency of the new approach.
TL;DR: In this article, an analytical approach is developed for investigating thermally induced vibrations of a split-blanket solar array due to self-shadowing of the central truss of the Space Station Freedom (SSF) design.
Abstract: An analytical approach is developed for investigating thermally induced vibrations of a split-blanket solar array due to self-shadowing of the central truss. Two analyses are developed: (1) cross-member shadowing of the truss longerons due to a torsional vibration of the solar array, and (2) parallelmember shadowing when the solar vector is aligned with the plane of the truss longerons. The analytical approaches identify key parameters for understanding the thermalstructural response. For parallel-member shadowing a stability analysis establishes the condition for thermal flutter. Computations are made for a solar array representative of the Space Station Freedom (SSF) design. The results show that cross-member shadowing is unlikely to cause thermally induced vibrations. Parallel-member shadowing can cause thermally induced vibrations; however, the vibrations are stable. Under normal operations the SSF solar array should not experience thermally induced vibrations.
TL;DR: In this article, an analysis of laminar hypersonic near-wake flows is presented, where the influence of nose bluntness and Mach and Reynolds numbers on the flowfield has been investigated, and the effects on the structure of the recirculation wake have been determined.
Abstract: An analysis of laminar hypersonic near-wake flows is presented. The influence of nose bluntness and Mach and Reynolds numbers on the flowfield has been investigated, and the effects on the structure of the recirculation wake have been determined. Analytical correlations for characterizing the near-wake behavior with regard to compressibility effects, extent of upstream influence, separation location, and near-wake aspect ratio have been established by means of numerical simulations and theoretical considerations. In particular the properties of the near wake are found to depend on the local Reynolds number along the recirculation dividing streamline: the base flow changes from a diffusionto a convection-dominated wake structure depending on the maximum value of the dividing-streamline Reynolds number. For a nearly inviscid recirculation a supersonic reverse flow occurs, with the formation of a secondary recirculation and a reverse-flow standing shock in the vicinity of the rear stagnation point. The upstream influence and the separation location have been shown to depend on the thickness of the subsonic boundary layer as well as the dividing streamline Reynolds number. Some of the conclusions have been validated by comparison with ground-test experimental data.
TL;DR: In this article, an eight-channel receiver is used to control the functions of the flight vehicle and servomechanisms are used as the actuating mechanisms, allowing only one set of signals to reach the flight controls, either from the autopilot or from the ground controller.
Abstract: Electrical System An eight-channel receiver is used to control the functions of the flight vehicle. Of the eight channels, four are assigned to the autopilot and primary flight controls. Servomechanisms are used as the actuating mechanisms. During autopilot operation the three primary flight control channels (elevator, rudder, and ailerons) from the receiver are disabled, allowing signals from only the autopilot to activate the servomechanisms. Conversely, when the autopilot is disengaged, the three flight control channels are active. This approach allows only one set of signals to reach the flight controls, either from the autopilot or from the ground controller. The remaining four channels are assigned to engine start, wing lock release, landing-gear release, and flaps. At the instant of engine start, a timer is activated, and the wing lock is released at a predetermined elapsed time. This serves as a backup feature, allowing wing deployment to occur in the event that no signals are received from the ground controller. Trajectory Analysis An analysis of the trajectory was conducted with SORT.2 This program allows three-dimensional trajectory optimization. Only pointmass trajectories were analyzed. Engine burnout occurs 36 s after liftoff, at an altitude of 12,965 ft. The maximum altitude achieved by the vehicle is 18,870 ft. Concluding Remarks This paper indicates that it may be possible to design a simple reusable rocket. The next step will be to build and test such a vehicle. This design could be extended and adapted to larger vehicles in order to attain the higher altitudes that are required in some of the applications of sounding rockets, such as upper-atmosph eric experiments.
TL;DR: A methodology for determining the near-optimal operation of the propulsion system of hybrid air-breathing launch vehicles is derived, based on selecting propulsion-system modes and parameters that maximize a certain performance function from consideration of the energy-state model of the aircraft equations of motion.
Abstract: A methodology for determining the near-optimal operation of the propulsion system of hybrid air-breathing launch vehicles is derived. The method is based on selecting propulsion-system modes and parameters that maximize a certain performance function. This function is derived from consideration of the energy-state model of the aircraft equations of motion. The vehicle model reflects the many interactions and complexities of the multimode air-breathing and rocket engine systems proposed for launch-vehicle use. The method is used to investigate the optimal throttle switching of air-breathing and rocket engine modes, and to investigate the desirability of using liquid-oxygen augmentation in air-breathing engine cycles, the oxygen either carried from takeoff or collected in flight. It is found that the air-breathing engine is always at full throttle, and that the rocket is on full at takeoff and at very high Mach numbers, but off otherwise. Augmentation of the air-breathing engine with stored liquid oxygen is beneficial, but only marginally so.
TL;DR: In this paper, the effects of member failure on the dynamic response of a truss-type space structure were investigated. But the authors focused on the effects on the structural response of the member failure in terms of deformations and stresses.
Abstract: Besides the usual causes of structural failure such as material defects and fabrication and construction errors, damage to space structures can come from such adverse events as impact by foreign objects, docking, drastic temperature changes, and exposure to radiation and charged particles. This paper presents a methodology to determine effects of member failure on the dynamic response of a truss-type space structure. Emphasis has been given to effects of the dynamic nature of member failure on the structural response. Two types of member failure are considered: one, the sudden brittle-type damage or failure of a member, and the other, member snap or dynamic jump due to buckling. The study is specially directed toward potential progressive member failure in the structure. It includes the postbuckling regime of member behavior. A three-dimensional truss-type structure is analyzed. Results are presented to delineate dynamic effects of member failure on the overall structural response measured in terms of deformations and stresses.
TL;DR: In this paper, the authors used the versatile Chimera numerical technique to determine the aerodynamic properties of small cylindrical segments being ejected into the wake of a flared projectile, and compared the unsteady Chimera results with the quasi-static approach.
Abstract: : This report describes the application of the versatile Chimera numerical technique to a time-dependent, multibody projectile configuration. A computational study was performed to determine the aerodynamics of small cylindrical segments being ejected into the wake of a flared projectile. The complexity and uniqueness of this problem results from the segments being in relative motion to each other, embedded in a nonuniform wake flow, and requiring a time-dependent solution. Flow field computations for this multibody problem have been performed for supersonic conditions. The predicted flow field over the segments was found to undergo significant changes as the segments separated from the parent projectile. Comparison of the unsteady Chimera results with the quasi-static approach shows the difference in drag history to be significant which indicates the need for time-dependent solution techniques. A subsequent experimental program was conducted in the Army Research Laboratory's (ARL) transonic Range and the computed segment positions and velocities were found to be in good agreement with the experimental data. Aerodynamics, Multiple bodies, Unsteady flow, Chimera, Wake, Drag