Showing papers in "Journal of Spacecraft and Rockets in 2002"

Review and synthesis of roughness-dominated transition correlations for reentry applications

Abstract: Nomenclature a = constant; Fig. 1 C , C 0 = constants k = roughness element height, ft N k = average roughness element height, ft L = vehicle length, ft M = Mach number n = exponent; Fig. 1 N R = Poll’s transition parameter; Eq. (1) Reke = roughnessReynolds number based on height k and edge conditions Rekk = roughnessReynolds number based on height k and conditions at k Reμ = Reynolds number based on height μ and edge conditions U = velocity component parallel to test surface or velocity component perpendicularto attachment line, ft /s V = velocity component parallel to attachment line, ft/s X = generalized disturbanceparameter or axial coordinate along windward centerline x = coordinate perpendicular to attachment line, ft Y = generalized transition parameter ® = angle of attack, deg ± = smooth-wall laminar boundary-layer thickness, ft = Poll’s length scale [Eq. (2)], ft μ = smooth-wall laminar boundary-layermomentum thickness, ft 1 = viscosity, lbm/ft ¢ s o = kinematic viscosity, ft2/s 1⁄2 = density, lbm/ft

Deployable Tensegrity Reflectors for Small Satellites

Abstract: Future small satellite missions require low-cost, precision reflector structures with large aperture that can be packaged in a small envelope. Existing furlable reflectors form a compact package which — although narrow— is too tall for many applications. An alternative approach is proposed, consisting of a deployable ''tensegrity'' prism forming a ring structure that deploys two identical cable nets (front and rear nets) interconnected by tension ties; the reflecting mesh is attached to the front net. The geometric configuration of the structure has been optimized to reduce the compression in the struts of the tensegrity prism. A small-scale physical model has been constructed to demonstrate the proposed concept. A preliminary design o fa3m diameter, 10 GHz reflector with a focal length to diameter ratio of 0.4 that can be packaged within an envelope of 0.1 × 0.2 × 0.8 m 3 is presented. Nomenclature a = radius of tube cross section A = area b = number of bars D = aperture diameter E = Young's Modulus F = focal length H = depth j = number of joints L = triangle side length Le = effective length m = number of independent mechanisms N = mesh tension per unit length r = radius of gyration R = radius of approximating sphere s = number of independent states of self-stress T tension δrms = root-mean-square surface error θ = relative rotation between nets ρ = density

Electroactive polymers as artificial muscles: A review

Abstract: Electroactive polymers (EAP) are an emerging class of functional materials that respond to electrical stimulation with large displacement. This attractive characteristic earned them the name artificial muscles. Even though the actuation force and robustness of existing EAP materials require further improvement, there has already been a series of reported successes. The successful applications that were demonstrated include catheter steering element, miniature manipulator, dust wiper, miniature robotic arm, and grippers. Some of the currently considered applications may be difficult to accomplish, and it is important to examine the requirements to the level that current materials can address. Using EAP to replace existing actuators may be a difficult challenge, and therefore it is highly desirable to identify a niche application where it would not need to compete with existing capabilities. The field involves multidisciplines that include materials, chemistry, electromechanics, computers, electronics, etc. A review of the state of the art and some of the challenges to the application of these materials are provided.

Graphical Method for Gravity-Assist Trajectory Design

Abstract: A new analytical technique directly related to Tisserand's criterion that permits the quick identification of all viable ballistic gravity-assist sequences to a given destination is introduced. The method is best presented by a simple graphical technique. The graphical technique readily demonstrates that gravity assists via Venus, Earth, and Jupiter are tremendously effective in sequences such as Venus-Earth-Earth-Jupiter and Venus-Earth-Mars-Earth. Estimates are made for the shortest flight times for a given launch energy to each planet. This graphical technique should provide mission designers with a potent tool for finding economical gravity-assist trajectories to many targets of high scientific interest in the solar system.

Design study of a square solar sail architecture

Abstract: The impact on the achievable performance and some key structural characteristics in 100 m square solar sails of design conditions and parameters is explored. Upper bound performance is addressed with an architecture of manageable mechanics and likely ultimate structural efficiency in its class due to uniaxial tensioning in each quadrant: the stripped sail. The designs are performed with a limit point-friendly approach. These innovations are justified by reviewing response fundamentals including sail billow, boom mechanics, and operational failure modes. The observations and conclusions advance the state of the art in sail design and provide guidance for future engineering. ac bj L, E F p Reu r, d t, S, s 77 p

Optimal spacecraft rendezvous using genetic algorithms

Abstract: The optimal rendezvous of two spacecraft are examined using a genetic algorithm. The minimum fuel solution of the optimal rendezvous contains many local optima, as well as discontinuities in the solution. These local optima and discontinuities make locating a global optimal solution difficult. Genetic algorithms are effective in solving these kinds of problems. The goal is to find the thrust time history that includes the magnitude and direction of the velocity change and the burn position, such that the boundary conditions are satisfied to an acceptable level and in a reasonable time. In addition, the number of thrust arcs and the maximum magnitude of the velocity change are regulated. This method was used on three test cases: 1) the Hohmann transfer, 2) the bielliptic transfer, and 3) rendezvous with two impulses. The results of the Hohmann and the bielliptic transfers match the analytical solutions within the resolution of the variables of the genetic algorithm. Although the result from the rendezvous with two impulses is not exact, the configuration of the trajectory is similar to the analytical solution.

Design of the First Interplanetary Solar Electric Propulsion Mission

Abstract: Deep Space 1 was the e rst interplanetary mission to be propelled by solar electric propulsion. The detailed design, development, and analysis of its trajectory have led to important new insights into the design of low-thrust trajectories. Tying the testing of solar electric propulsion technology to an operational mission has allowed the identie cation of trajectory design issues that were not considered in concept studies, such as constraints on the spacecraft attitude and periods of thrusting or coasting that are dictated by reasons other than trajectory considerations. Models of the spacecraft performance unimportant for trajectory analysisin missions using conventional chemical propulsion are intimately connected with the design of the trajectory when solar electric propulsion is employed.In addition, massmargin is notsufe cientto assessa low-thrustmission. Unplanned thrust interruptions may result in a situation in which the propulsion system cannot provide sufe cient impulse to compensate for the lost thrust in time to reach encounter targets. Mission margin (as distinct from mass margin ) is a quantie cation of the mission’ s susceptibility to loss of thrust and is an important indicator of mission robustness for low-thrust trajectories.

Space Systems Flexibility Provided by On-Orbit Servicing: Part 1

Abstract: Spacecraft are still the only complex engineering systems without routine maintenance infrastructure. Whereas the technologies for autonomous on-orbit servicing of satellites are emerging, no general conclusions have yet been drawn regarding the cost effectiveness of on-orbit servicing. In a companion paper (Saleh, J. H., Lamassoure, E., and Hastings, D. E., "Space Systems Flexibility Provided by On-Orbit Servicing: Part 1," Journal of Spacecraft and Rockets, Vol. 39, No. 4, 2002, pp. 551-560), a new perspective on the problem was proposed, in which the value of servicing is studied independently from its cost. A framework to account for the value of the flexibility provided by on-orbit servicing to space systems was developed. Here, the usefulness of this framework is demonstrated by studying the value of servicing for two types of space missions. First commercial missions with uncertain revenues are considered. It is shown that traditional valuation has been underestimating mission value by not taking into account the option to abandon. Then servicing is considered as an option on life extension, showing how the optimal design life decreases with increasing uncertainty. A map of the maximum servicing price in a market level/market volatility space is proposed as a new tool for decision making. Then military missions faced with uncertainty in the location of contingencies are considered. The value of refueling for making spacecraft maneuverable is studied for two cases. For a radar constellation in low Earth orbit, servicing is shown to have little value due to a conflict between propulsion mass and maneuver time. For a geostationary fleet of communication satellites, servicing is shown to have value based on the potential improvements in capacity usage.

Hypersonic Flow over a Blunt Body with Plasma Injection

Abstract: The drag reduction of blunt body in hypersonic flow via plasma injection has been investigated by a combined experimental and computational effort. The counterflow plasma jet generated by a plasma torch has a vibrionic temperature of 4400 K, an electronic temperature around 20,000 K, and electron density greater than 3 × 1012/cm3. At a fixed injection stagnation pressure and in the absence of an applied magnetic field, the plasma injection actually increases drag above that of room-temperature air due to a decreased mass flow rate at the elevated temperature. However, at an identical mass flow rate, the plasma injection reveals a greater drag reduction than the room-temperature air counterpart through thermal energy deposition. From experimental measurements, an overwhelming major portion of the drag reduction is derived from the viscous-inviscid interaction of the counterflow jet and thermal energy deposition. The numerical results of Navier-Stokes equations with a local equilibrium plasma composition also confirm this observation.

Six-Axis Vibration Isolation System Using Soft Actuators and Multiple Sensors

Abstract: Several types of Stewart platforms have been implemented by research groups to examine design and control issues in six-axis vibration isolation for space-based systems. Hood Technology Corporation and the University of Washington have taken the lessons learned from these various designs and developed a new hexapod that addresses the requirements of the Jet Propulsion Laboratory's planned spaceborne interferometry missions. This system is unique in its very soft axial stiffness (3-Hz corner frequency) for active isolation and pointing control, custom-designed voice coil actuator with a large displacement capability, and elastomeric flexures both for guiding the actuator and providing pivots at the end of each strut. In addition, there are four sensors in each strut for control topology design and evaluation. An overview of this unique six-axis isolator design and a summary of the control results for various sensor topologies, including multisensor and frequency-weighted isolation and pointing control, are presented. Controllers that experimentally achieved 20-25-dB reduction in vibration in all six degrees of freedom across the bandwidth of interest (5-20 Hz) are shown.

Analysis of Bare-Tether Systems for Deorbiting Low-Earth-Orbit Satellites

Abstract: Performances, design criteria, and system mass of bare tethers for satellite deorbiting missions are analyzed. Orbital conditionsand tethercrosssection dee nea tetherlength, suchthat1 )shortertethers areelectron collecting practically in their whole extension and 2 ) longer tethers collect practically the short-circuit current in a e xed segment length. Long tethers have a higher drag efe ciency (dee ned as the drag force vs the tether mass ) and are better adapted to adverse plasma densities. Dragging efe ciency and mission-related costs are used to dee ne design criteria for tether geometry. A comparative analysis with electric thrusters shows that bare tethers have much lower costs for low- and midinclination orbits and remain an attractive option up to 70 deg.

In-Flight Estimation of the Cassini Spacecraft's Inertia Tensor

Abstract: A new methodology for estimating the spacecraft 3-by-3 inertia tensor is proposed in this study.

Corner Wrinkling of a Square Membrane Due to Symmetric Mechanical Loads

Abstract: Thin-film membrane structures are under consideration for use in many future gossamer spacecraft systems. Examples include sunshields for large aperture telescopes, solar sails, and membrane optics. The development of capabilities for testing and analyzing pre-tensioned, thin film membrane structures is an important and challenging aspect of gossamer spacecraft technology development. This paper presents results from experimental and computational studies performed to characterize the wrinkling behavior of thin-fi[m membranes under mechanical loading. The test article is a 500 mm square membrane subjected to symmetric comer loads. Data is presented for loads ranging from 0.49 N to 4.91 N. The experimental results show that as the load increases the number of wrinkles increases, while the wrinkle amplitude decreases. The computational model uses a finite element implementation of Stein-Hedgepeth membrane wrinkling theory to predict the behavior of the membrane. Comparisons were made with experimental results for the wrinkle angle and wrinkled region. There was reasonably good agreement between the measured wrinkle angle and the predicted directions of the major principle stresses. The shape of the wrinkle region predicted by the finite element model matches that observed in the experiments; however, the size of the predicted region is smaller than that determined in the experiments.

Automated Design of the Europa Orbiter Tour

Abstract: Before the Europa Orbiter can be placed in orbit about Europa, it will be placed into a 200-day Jovian orbit and targeted for Ganymede. After a series of gravity-assist flybys of the Galilean satellites, the orbital energy is reduced to lower the arrival hyperbolic excess velocity at Europa. These energy-saving techniques reduce the propellant cost for Europa orbit insertion to a minimum. Key constraints during the tour include total time of flight and radiation dosage. Tours may employ 10 or more encounters with the Jovian satellites; hence, there is an enormous number of possible sequences of these satellites. A graphical method based on Tisserand's criterion is presented that greatly aids the design process. The Tisserand graph method facilitates the study of a wide range of arrival conditions, arrival dates, and satellite tours.

Thermal Analysis and Design Optimization of Multilayer Insulation for Reentry Aerodynamic Heating

Abstract: A design-of-experiment technique was used to investigate optimum design of multilayer insulations for reentry aerodynamic heating. Combined radiation/conduction heat transfer in high-temperature multilayer insulations wasmodeled using a finite volume numerical model, which was validated by comparison with steady-state effective thermal conductivity measurements, and transient thermal tests simulating reentry aerodynamic heating conditions. It was found that use of 2-mm foil spacing and locating the foils near the hot boundary, with the top foil 2 mm away from the hot boundary, resulted in the most effective insulation design. A 76.2-mm-thick multilayer insulation using 1, 4, or 16 foils resulted in 2.9, 7.2, or 22.2% mass per unit area savings, respectively, compared to a fibrous insulation sample at the same thickness.

Spacecraft Design Lifetime

Abstract: A general discussion of issues that drive and limit spacecraft design lifetime is presented. The effects of varying the spacecraft lifetime requirement on different subsystems are explored, and typical spacecraft mass andcost profiles are deduced. Quantitative analyses confirm that the design lifetime is a key requirement in sizing various subsystems and significantly affects the spacecraft mass and cost to initial operating capability. The analysis introduces a formally defined economic metric, the cost per operational day, to help guide the specification of the design lifetime requirement. Preliminary results suggest that other factors should also be taken into account in specifying the design lifetime, namely, the loss of value resulting from technology obsolescence as well as the volatility of the market the system is serving in the case of a commercial satellite.

Supersonic/Hypersonic Flutter and Postflutter of Geometrically Imperfect Circular Cylindrical Panels

Abstract: A theoretical investigation of the flutter and postflutter of infinitely long thin-walled circular cylindrical panels in a supersonic/hypersonic flowfield is presented. In this context, third-order piston theory and shockwave aerodynamics are used in conjunction with the geometrically nonlinear shell theory to obtain the pertinent aeroelastic governing equations. The effects of in-plane edge restraints and small initial geometric imperfections are also considered in the model. The objective is twofold: 1) to analyze the implications of nonlinear unsteady aerodynamics and structural nonlinearities on the character of the flutter instability boundary and 2) to outline the effects played, in the same respect, by a number of important geometrical, physical, and aerodynamic parameters characterizing the aeroelastic system. As a by-product of this analysis, the implications of these parameters on the linearized flutter instability behavior of the system are captured and emphasized. The behavior of the aeroelastic system in the vicinity of the flutter boundary is studied via the use of an encompassing methodology based on the Lyapunov first quantity. Numerical illustrations, supplying pertinent information on the implications of geometric and aerodynamic nonlinearities, as well as of other effects, such as curvature and thickness ratios, on the flutter instability and on the character of the flutter boundary are examined, and pertinent conclusions are outlined.

Numerical Investigation of Supersonic Jet Interaction for Finned Bodies

Abstract: : A detailed numerical investigation of the interaction between a lateral jet and the external flow has been performed for a variety of missile body geometries. These include non-finned axisymmetrical bodies and finned bodies with either strakes or aft-mounted tail fins. The computations were performed at Mach numbers 2, 4.5, and 8. To obtain the numerical results, both Reynolds- averaged Navier-Stokes and Euler techniques were applied. The computational results were compared with results from a previously published wind tunnel study that consisted primarily of global force and moment measurements. The results show significant interactions of the jet-induced flow field with the fin surfaces, which produce additional effects compared with the body alone. In agreement with the wind tunnel study, in some cases the presence of lifting surfaces resulted in force and/or moment amplification of the jet interaction with the missile surfaces. The results indicate deamplification of the jet force at Mach 2 for all three bodies. Amplification of the jet force was also observed for high Mach numbers, particularly for the body with strakes. For the results examined here, there were only minor differences in the global force and moment predictions when viscous or inviscid techniques were used. The dependence of the interaction parameters on angle of attack and jet pressure was well predicted by both methods. The numerical techniques showed good agreement with the experiments at supersonic Mach numbers but only fair agreement for the hypersonic, Mach 8 case.

Applications of a Counterflow Drag Reduction Technique in High-Speed Systems

Abstract: Potential applications of a counterflow drag-reduction technique were investigated to assess performance improvements on aerospace vehicles. The motivation for this study was the 30-50% drag reduction achieved by counterflow blowing experiments on hemispherical cylinders at Mach 4 and higher. Exploratory studies indicate that drag improvements by counterflow drag reduction on hemispherical bodies cannot match those of aerodynamically shaped sharp-nosed bodies. Hence, the approach taken in the present study is that for hypersonic Mach numbers: if the nose shape is required to be blunt for considerations other than drag, counterflow blowing can be effective in improving the performance of the system. Although for generic body shapes counterflow blowing is most effective for blunt-nosed bodies, when applied to actual systems, many other factors need to be considered, such as available internal volume and extreme compressed carriage requirements. Depending on the vehicle speed and nose shape, estimated drag reductions of 15-30% were applied to predict the overall performance gains on Space Operations Vehicle, Gun-Launched Rocket, and Pegasus XL configurations. Potential savings in propellant and improvements in burnout velocity and range are reported. For launch systems with high fuel fraction, the payoff with counterflow drag reduction is marginal as the overall effects of aerodynamic drag on performance are small in the upper atmosphere. For the lower fuel fraction vehicle, the Gun Launched Rocket, a range improvement of 7% was achieved for a drag reduction of 30% with 0.3 blunting of nose flying above Mach 3; with greater blunting, however, the volume of fuel cannot compensate for the increase in drag.

Plume Modeling of Stationary Plasma Thrusters and Interactions with the Express-A Spacecraft

Abstract: Hall-effect thruster flight measurements are compared with results from two-dimensional plume and three-dimensional spacecraft interactions computer simulations. The measurements were acquired onboard Express-A 2 and A 3,two Russian communications satellites in geosynchronus orbit. The spacecraft carry four propulsion units for east-west and north-south station keeping. Each unit consists of two stationary plasma thrusters. Ion flux and energy spectra were recorded at various positions with respect to the thrusters and are compared with results from simulations using a uniform electron temperature, two-dimensional plume code that computes the expansion of the main ion beam by a fluid approach. The dynamics of the charge-exchange plasma are determined by a particle-in-cell method. Comparisons suggest good agreement for plume angles less than 40 deg and electron temperature between 8 and 11 eV. At approximately 4 and 9 m away from the thruster, and at plume angles less than 10 deg, the discrepancy between measured and computed values is found to be less than 10%. At larger angles, ion flux measurements exhibit large variations during operation of the same thruster. At 80 deg and 1.35 m away from the thruster, flux sensors recorded current densities that ranged between 12 and 55 mA/m 2 . The two-dimensional code computes 27 mA/m 2 for an anode mass flow rate of 5.3 mg/s at this location. Moments induced on the spacecraft during the operation of each thruster were also recorded by the attitude control system and are compared with results from a three-dimensional spacecraft interactions code. These measurements were taken during rotation of the solar arrays.

Rationale for Defining Structural Requirements for Large Space Telescopes

Abstract: The present paper presents a rationale for defining structural requirements for future large space telescope systems. The rationale is based on bounding analyses for the deformation of telescope mirrors in response to expected on-orbit disturbance loads and consideration of active control systems that partially compensate for these deformations. It is shown that the vibration frequency of the telescope structure, independent of telescope size, determines the passive structural stability and requirements for an active control system. This means that future large telescopes with low vibration frequencies will necessarily allocate increased active control error budget in proportion to the square of the vibration frequency. Parametric analyses are also presented for the vibration response of two representative mirror architectures: a tensioned membrane mirror and a truss-supporte d segmented mirror. These examples demonstrate that meeting a specified frequency requirement will require a trade between structural mass fraction and depth of the primary mirror support structure regardless of the structural architecture.

Plume Study of a 1.35-kW SPT-100 Using an ExB Probe

Abstract: The ion energy distributions of Xe 1+ , Xe 2+ , and Xe 3+ ions in the SPT-100 plume obtained 50 cm and 1 m from the thruster exit with an ExB probe are presented. Most of the ion species distribution functions exhibit features associatedwithbothMaxwellianandDruyvesteyndistributions.Therefore,thenatureoftheionenergydistribution in the SPT-100 plume is established by the competing effects of ion acceleration in the discharge chamber and collisional processes beyond the ion production zone. Comparison of beam energy and ion energy spread 50 cm and 1 m from the thruster exit reveal that the energy distribution of the far-e eld plume ions varies little as the ions move away from the thruster. The angular proe les of the ion species fractions and the beam energy data suggest that, whereas Xe 2+ and Xe 3+ ions are produced near the thruster exit, Xe 1+ ions are created farther upstream in

Experimental and Numerical Investigation of Martian Atmosphere Entry

Abstract: A numerical and experimental investigation was performed to study the aerothermodynamic problems of entry into the Martian atmosphere. The mathematical and physical model used to study the flowfield around a capsule entering a CO 2 environment is described. Computational fluid dynamics tools have been applied to solve the system of governing equations. The importance of surface catalycity effects on the stagnation-point heat transfer and on the heat load in Martian atmosphere is highlighted. Stagnation-point heat flux levels applied to models of different materials in a plasma wind tunnel are shown, and numerical correlations are presented. The different role played by surface catalycity in Earth and Mars environments is shown.

Modeling of the interaction of a side jet with a rarefied atmosphere

Abstract: The interaction of a jet from a 3000-N-class thruster positioned on the side of a small rocket, with the rarefied atmosphere at 100 and 80 km, is studied numerically. The direct simulation Monte Carlo method was applied to model the three-dimensional jet-atmosphere interaction. Chemical reactions between freestream and plume species were included in the simulations. A two-stage numerical strategy was used, with sequential computations of an axisymmetric plume core flow and three-dimensional plume-freestream interaction. The impact of altitude, angle of attack, rocket velocity, and thrust on flowfields and surface mass fluxes is examined.

Design of aerogravity-assist trajectories

Abstract: Aero-gravity assist (AGA) trajectories are optimized in the sense of maximizing AV obtained by the fly by, maximizing aphelion, minimizing perihelion, and minimizing the time of flight (TOF) for a particular destination planet. A graphical method based on Tisserand's criterion is introduced to identify potential AGA trajectories. To demonstrate the application of the theory, patched-conic AGA trajectories are computed to each planet in the Solar System. For an L/D of 7, and a launch VQO of 6.0 km/s, Pluto may be reached in 5.5 years using a Venus-Mars-Venus AGA.

Dimensional Repeatability of an Elastically Folded Composite Hinge for Deployed Spacecraft Optics

Abstract: A new type of folded composite hinge is investigated for its use in precision deployable spacecraft structures. The hinge is an integral feature of a composite tube intended for use as a structural truss member. The design of the hinge allows the tube to be elastically folded for stowage even with tube wall thicknesses from 0.4 to 1.7 mm. Whether the large but primarily elastic folding stresses impart permanent deformations to the tube after it is deployed is experimentally assessed. The data show that any such permanent strain induces tip deformations, identified as microscopic plastic behavior, of no more than 2.5 μ axially and 9 p laterally, depending on the composite layup. This deployment repeatability is comparable to prior measurements of mechanical deployables. Moreover, stow duration and number of stows have no measurable effect, once the initial stow-deploy cycle has been completed. There is always a significant viscoelastic creep recovery following deployment that increases with stowage time. However, this viscoelastic creep is recovered. An exponential curve fit of the creep time response shows that the time constants of the viscoelastic recovery are independent of stow duration.

Realizing "2001: A Space Odyssey": Piloted Spherical Torus Nuclear Fusion Propulsion

Abstract: A conceptual vehicle design enabling fast, piloted outer solar system travel was created predicated on a small aspect ratio spherical torus nuclear fusion reactor. The initial requirements were satisfied by the vehicle concept, which could deliver a 172 mt crew payload from Earth to Jupiter rendezvous in 118 days, with an initial mass in low Earth orbit of 1,690 mt. Engineering conceptual design, analysis, and assessment was performed on all major systems including artificial gravity payload, central truss, nuclear fusion reactor, power conversion, magnetic nozzle, fast wave plasma heating, tankage, fuel pellet injector, startup/re-start fission reactor and battery bank, refrigeration, reaction control, communications, mission design, and space operations. Detailed fusion reactor design included analysis of plasma characteristics, power balance/utilization, first wall, toroidal field coils, heat transfer, and neutron/x-ray radiation. Technical comparisons are made between the vehicle concept and the interplanetary spacecraft depicted in the motion picture 2001: A Space Odyssey.

Particle-in-Cell with Monte Carlo Simulation of SPT-100 Exhaust Plumes

Abstract: A two-dimensional axisymmetric numerical code is developed for the simulation of a Hall thruster plume operating in various ambient plasmas. The code is based on a combination of particle simulation for the ionic components (Xe + and Xe + + ) and fluid computational techniques for electrons. In particular, we have used the Boltzmann relation modified in order to allow for the effect of nonisothermal electron temperature based on the adiabatic approximation. In our model several solutions, which have been sparsely considered in previous works, are jointly adopted. In particular, the electric field is computed by solving the Poisson equation without assuming quasi-neutrality, which often is violated in the near-field plume region; and collision processes are included by using two new techniques, the ion-neutral test-particle Monte Carlo collision model and the Nanbu cumulative small-angle collision theory for ion-ion coulombic collisions. Comparisons with experimental data suggest that the present simulation is accurately modeling the physics of the very near-field region of the plume.

Design of a Hybrid Chemical/Electric Propulsion Orbital Transfer Vehicle

Abstract: A hybrid chemical/electric propulsion orbital transfer vehicle was studied for a coplanar transfer from low Earth orbit to geosynchronous orbit. The goal of this study was to develop a strategy combining the two propulsion systems to reduce total radiation loads on the spacecraft and to provide attractive total trip times. Propulsion systems were sized based on current technology storable bipropellants for the chemical stage; three separate electric propulsion systems were considered for the low-thrust portion of the mission. The effect of the relative amounts of chemical propulsion on trip time, vehicle mass, and overall radiation dosage has been characterized. In addition, the effect of orbit inclination, the altitude for initiation of the chemical propulsive burns, and payload mass have been considered. Attractive options have been identified that are substantially lighter than a pure chemical system and that offer trip times and radiation loads substantially less than a pure electric system.

Everything Is Relative in Spacecraft System Alignment Calibration

Abstract: The concepts of absolute and relative alignment calibration of spacecraft attitude sensors are examined. It is known that three degrees of freedom of attitude associated with absolute alignment calibration are unobservable unless payload data are processed. It is shown that an absolute alignment model is equivalent to a relative alignment model when the payload is regarded as an attitude sensor. Then it is shown that the three unobservable degrees of freedom are eliminated by defining the gyro as the body reference frame, attributing only three nonorthogonal misalignment parameters to the gyro. The payload misalignment can then be parameterized and calibrated in the same manner as any attitude sensor, or this can be left strictly to the payload data processing, thus creating a well-defined boundary between attitude control system calibration and payload calibration. The new parameterization introduced is illustrated via simulation results.