Showing papers in "Journal of Spacecraft and Rockets in 2006"

History of Space Shuttle Rendezvous and Proximity Operations

Abstract: Space Shuttle rendezvous missions present unique challenges that were not fully recognized when the Shuttle was designed Rendezvous targets could be passive (ie, no lights or transponders), and not designed to facilitate Shuttle rendezvous, proximity operations, and retrieval Shuttle reaction control system jet plume impingement on target spacecraft presented induced dynamics, structural loading, and contamination concerns These issues, along with limited reaction control system propellant in the Shuttle nose, drove a change from the legacy Gemini/Apollo coelliptic profile to a stable orbit profile, and the development of new proximity operations techniques Multiple scientific and on-orbit servicing missions, and crew exchange, assembly and replenishment flights to Mir and to the International Space Station drove further profile and piloting technique changes These changes included new proximity operations, relative navigation sensors, and new computer generated piloting cues However, the Shuttle's baseline rendezvous navigation system has not required modification to place the Shuttle at the proximity operations initiation point for all rendezvous missions flown

Review on Dynamics and Control of Nonelectrodynamic Tethered Satellite Systems

Abstract: L= distance between mass centers of satellites m1 and m2, tether length if m1 � m2 ,m L 0 = initial tether length, m M = total system mass, kg mi = mass of satellite i ,k g mt = tether mass, kg ri = position vector of mass mi from the system center of mass, m ˆ r1 = r1/a β, η = relative in-plane and out-of-plane swing angles between m1 and m2, respectively, deg βmax ,η max = maximum values of β and η, respectively, deg ˙ β0 = ˙ β at θ = 0 � Ha1 = maximum altitude decrease of m1 at apogee, m � Hi = altitude gain of satellite mass mi ,m � Hp2 = maximum altitude gain of m2 at perigee, m � L = decrease in the initial tether length L 0 during tether retrieval, m δ = ratio of in-plane swing rate ˙ β to the orbital rate � μ = gravitational constant, m 3 s −2 � = orbital rate, (μ/a 3 ) 1/2 , rad/s

Preliminary Design of Low-Thrust Multiple Gravity-Assist Trajectories

Abstract: The availability of electric engines as primary sources of propulsion has opened the doors to new scenarios for future interplanetary missions, but has increased the complexity of trajectory design. This paper proposes a novel approach to the preliminary design of interplanetary trajectories characterized by a combination of low-thrust propulsion and multiple gravity-assist maneuvers. Low-thrust arcs are obtained by shaping the trajectory through a set of parameterized pseudoequinoctial elements. The characterization of the solution space for a particular set of planetary encounters and a range of launch dates is then performed through a global optimization method, blending a particular evolutionary algorithm with a deterministic domain decomposition technique. The effectiveness of the proposed approach is demonstrated through a number of examples of the design of low-thrust, gravity-assist interplanetary trajectories.

Experiment of the Two-Dimensional Roughness Effect on Hypersonic Boundary-Layer Transition

Abstract: An experimental investigation of the effects of two-dimensional roughness on hypersonic boundary layers was carried out at the JAXA 0.5 m hypersonic wind tunnel using a 5 deg half-angle sharp cone at a freestream Mach number of 7.1 and a wide range of stagnation conditions. Aerodynamic heating distributions and surface pressure fluctuations were measured with and without roughness elements applied to the cone. A wavy wall roughness with a wavelength of 2� located well upstream of the breakdown region had the effect of delaying transition. Surface pressure spectrum densities indicate that disturbances of the roughness wavelength are amplified compared to the smooth wall case. At a lower stagnation temperature, however, wavy wall roughness configurations had the same effect as regular roughness, and no discernible differences between wavy wall roughness and spherical roughness were observed, either in the transition point or in the pressure fluctuation spectrum.

Preliminary design of multiple gravity-assist trajectories

Abstract: In this paper the preliminary design of multiple gravity-assist trajectories is formulated as a global optimization problem. An analysis of the structure of the solution space reveals a strong multimodality, which is strictly dependent on the complexity of the model. On the other hand it is shown how an oversimplification could prevent finding potentially interesting solutions. A trajectory model, which represents a compromise between model completeness and optimization problem complexity is then presented. The exploration of the resulting solution space is performed through a novel global search approach, which hybridizes an evolutionary based algorithm with a systematic branching strategy. This approach allows an efficient exploration of complex solution domains by automatically balancing local convergence and global search. A number of difficult multiple gravity-assist trajectory design cases demonstrates the effectiveness of the proposed methodology.

Experimental Study on Thermal Protection System by Opposing Jet in Supersonic Flow

Abstract: Introduction C URRENTLY, developments of reusable launch vehicle (RLV) for a low-cost space transportation system are in progress. In the development of RLV, one of the most important problems is the severe aerodynamic heating at the nose and leading edges of the vehicle. In such supersonic and hypersonic flights, prediction of aerodynamic heating and construction of proper thermal protection system are especially important. Heat-resistant tiles and ablators are currently used for thermal protection systems. However, those thermal protection systems are not reusable. In the present study, the method using an opposing jet is proposed for fully reusable thermal protection system of RLV. The method can be considered to have almost the same effect of heat reduction at nose region as the method with mechanical spike.1 The opposing jet works as an aerodynamic spike to move the detached shock wave away from the nose and form a recirculation region, which is quite effective to reduce aerodynamic heating at the nose region. The schematic diagram of supersonic flowfields with opposing jet injected at the nose of a blunt body is shown in Fig. 1. In the flowfield, the opposing jet forms a Mach disk and contact surface with freestream. The jet layer reattaches to the body surface and forms a recirculation region between the nozzle exit and reattachment point of the jet layer. The recompression shock wave is formed near the reattachment point of the jet layer. Many studies on opposing jet flow have been conducted in order to reveal the flow mechanism.2−7 However, most of those studies are related to the stability of flowfield and oscillations of shock waves. Except for Warren,6 not much study has been conducted to reveal the effects of opposing jet on reduction of aerodynamic heating. In the present study, geometric ratio of diameters and Mach number are fixed. The flow stability is determined by the total pressure ratio of freestream to opposing jet. We define the total pressure ratio

Vibration Isolator for Large Space Telescopes

Abstract: A full spacecraft vibration isolator for the James Webb Space Telescope is described. This 1-Hz isolator brings wavefront errors and line-of-sight pointing jitter induced by reaction wheel and cryocooler compressor disturbances down to below a few nanometers and milliarcseconds, respectively. The isolator consists of four passively damped beams connecting the corners of the spacecraft to a thermal isolation tower positioning the telescope. An efficient analysis approach was developed based upon a fractional derivative model of the viscoelastic material. The technique employs modal data from a finite element model for a constant operating temperature and frequency. The modal properties are adjusted using frequency- and temperature-dependent properties from the viscoelastic material’s constitutive model. Development test results for the unit isolator elements are presented and compared to model predictions. Results of a dynamic test on an assembled isolator are also presented. The test employs a simulator replicating the full mass and inertia of the telescope. A series of modal surveys was performed across the operating temperature range to validate model predictions.

Mars Science Laboratory Simulations for Entry, Descent, and Landing

Abstract: Two primary simulations have been developed and are being updated for the Mars Science Laboratory entry, descent, and landing. The high-fidelity engineering end-to-end entry, descent, and landing simulation is based on NASA Langley Research Center's Program to Optimize Simulated Trajectories II and the end-to-end real-time, hardware-in-the-loop simulation test bed, which is based on NASA Jet Propulsion Laboratory's Dynamics Simulator for Entry, Descent, and Surface landing. The status of these Mars Science Laboratory entry, descent, and landing end-to-end simulations at this time is presented. Various models, capabilities, as well as validation and verification for these simulations, are discussed.

Laminar-Turbulent Transition on Reentry Capsules and Planetary Probes

Abstract: Introduction N ASA interest in reentry vehicles has recently taken a dramatic shift away from lifting bodies (X-33, X-38) and winged vehicles (shuttle, X-34) towards capsules. This follows the loss of the Columbia and the Presidential announcement of a program focused on the moon and Mars [1]. The Presidential Commission report mentions “entry, descent, and landing: precision targeting and landing on ‘high-g’ and ‘low-g’ planetary bodies” as one of 17 critical areas identified for the initial focus ([1], p. 28). NASA is developing a crew exploration vehicle (CEV) to fulfill these missions, following the planned retirement of the shuttle circa 2010. It has been determined that this CEV is to be an Apollo-like capsule. Although moderate levels of lift can much improve the safety and flexibility of entry trajectories ([2], Chap. 6), the Apollo geometry is proven in the very high aeroheating of a moon-return entry {11 km=s ([3], p. 19)}, and easier to integrate with a launchescape rocket. The sameCEV reentry geometry is to be used for both the low-energy entries from low earth orbit (LEO) and the very different high-energy entries from hyperbolic lunar and interplanetary missions. The planetary program also continues to fly blunt bodies as entry probes; it seems efficient to consider these similar shapes in parallel with the capsules. Transition is one aerothermodynamic problem that is not yet well understood. Will the uncertainties in estimating transition [4] impact the design of these vehicles? Is transition important to the development of new capsules a la Mercury, Gemini, Apollo, or Soyuz? It is generally thought that transition is more important for slender vehicles than for blunt vehicles, although few studies have been carried out to support this idea [5]. Because there is much uncertainty about the properties of any such capsules (geometry, ballistic coefficient, mission needs, and so on), this question cannot be answered definitively. Nevertheless, the following reviews the public-domain literature from blunt-body and capsule programs, with a view towards identifying the available data for transition under these conditions, the physics of the transition process, and the conditionswhere laminar-turbulent transition is important. Particular attention is paid to manned capsules, where funding levels are likely to support more detailed analyses. Although the Russian Soyuz vehicle may be an important forerunner to the CEV [6], detailed technical information was not available to the author, so it is not discussed. This survey, updated from [7], is certainly incomplete, and the author would appreciate any additional information that the reader might be able to provide. Laminar-turbulent transition in hypersonic boundary layers is important for prediction and control of heat transfer, skin friction, and other boundary layer properties. Vehicles that spend extended periods at hypersonic speeds may be critically affected by the uncertainties in transition prediction, depending on their trajectories, geometries, and surface properties. However, the mechanisms leading to transition are still poorly understood, even in low-noise environments. Many transition experiments have been carried out in conventional ground-testing facilities over the past 50 years [4]. However, these experiments are contaminated by the high levels of noise that radiate from the turbulent boundary layers normally present on the wind tunnel walls [8]. These noise levels, typically 0.5–1% of the mean, are an order of magnitude larger than those observed in flight [9,10]. These high noise levels can cause transition to occur an order ofmagnitude earlier than inflight [8,10]. In addition, themechanisms of transition operational in small-disturbance environments can be changed or bypassed altogether in high-noise environments; these changes in themechanisms change the parametric trends in transition [9]. Although transition can become completely dominated by roughness [11] or perhaps ablation effects, these effects are usually only one of several factors whose effect must be understood to reliably predict flight [9,12]. Because no single ground-test facility can simultaneously duplicate the velocity, scale, freestream noise, freestream chemistry, and surface temperature of reentry flight, partial simulations in the available facilities must be combined to develop computational models that can then be extrapolated toflight. Mechanism-based methods must be developed to provide reliable predictions. Such mechanism-based prediction methods are now becoming feasible for complex three-dimensional flows at hypervelocities, due to ever-increasing computational capabilities. For example, Johnson et al. recently provided stability-based transition analyses for a planetary probe, although their preliminary results are only for a simplified axisymmetric case [13]. Because the best mechanism-based methods will still require many assumptions, development and validation will require measurements of the

Design of the ARES Mars Airplane and Mission Architecture

Abstract: Significant technology advances have enabled planetary aircraft to be considered as viable science platforms. Such systems fill a unique planetary science measurement gap, that of regional-scale, near-surface observation, while providing a fresh perspective for potential discovery. Recent efforts have produced mature mission and flight system concepts, ready for flight project implementation. This paper summarizes the development of a Mars airplane mission architecture that balances science, implementation risk and cost. Airplane mission performance, flight system design and technology maturation are described. The design, analysis and testing completed demonstrates the readiness of this science platform for use in a Mars flight project.

Evaluation of Multidisciplinary Optimization Techniques Applied to a Reusable Launch Vehicle

Abstract: Three multilevel multidisciplinary optimization techniques, Bi-Level Integrated System Synthesis, Collaborative Optimization, and Modified Collaborative Optimization, are applied to the design of a reusable launch vehicle, evaluated, and compared in this study. In addition to comparing the techniques against each other, they are also compared with designs reached via fixed-point iteration of disciplines with local optimization and the industry accepted multidisciplinary optimization technique, All-at-Once. The new multidisciplinary optimization techniques, particularly Bi-Level Integrated System Synthesis, showed greater ability than fixed-point iteration to design for a global objective and were more applicable to complex systems than All-at-Once. This study was the first time that the novel multidisciplinary optimization methods were compared qualitatively and quantitatively under controlled experimentation practices. It is still impossible to statistically determine whether any one of the novel multidisciplinary optimization techniques is better than another, because more studies using different test problems corroborating the conclusions made here are needed.

Satellite Constellation Design Tradeoffs Using Multiple-Objective Evolutionary Computation

Abstract: Multiple-objective evolutionary computation provides the satellite constellation designer with an essential optimization tool due to the discontinuous, temporal, and/or nonlinear characteristics of the metrics that architectures are evaluated against. In this work, the nondominated sorting genetic algorithm 2 (NSGA-2) is used to generate sets of constellation designs (Pareto fronts) that show the tradeoff for two pairs of conflicting metrics. The first pair replicates a previously published sparse-coverage tradeoff to establish a baseline for tool development, whereas the second characterizes the conflict between temporal (revisit time) and spatial (image quality) resolution. A thorough parameter analysis is performed on the NSGA-2 for the constellation design problem so that the utility of the approach may be assessed and general guidelines for use established. The approximated Pareto fronts generated for each tradeoff are discussed, and the trends exhibited by the nondominated designs are revealed. Nomenclature a = semimajor axis, km e = eccentricity F = focal length, m i = inclination, deg K = units conversion constant M = mean anomaly, deg P = pixel size, μm e = elevation, deg ρ = range, km � = right ascension of the ascending node, deg ω = argument of perigee, deg

Solar Polar Orbiter: A Solar Sail Technology Reference Study

Abstract: An assessment is presented of a Solar Polar Orbiter mission as a Technology Reference Study. The goal is to focus the development of strategically important technologies of potential relevance to future science missions. The technology is solar sailing, and so the use of solar sail propulsion is, thus, defined a priori. The primary mission architecture utilizes maximum Soyuz Fregat 2-1b launch energy, deploying the sail shortly after Fregat separation. The 153 × 153 m square sail then spirals into a circular 0.48-astronomical-unit orbit, where the orbit inclination is raised to 90 deg with respect to the solar equator in just over 5 years. Both the solar sail and spacecraft technology requirements have been addressed. The sail requires advanced boom and new thin-film technology. The spacecraft requirements were found to be minimal because the spacecraft environment is relatively benign in comparison with other currently envisaged missions, such as the Solar Orbiter mission and BepiColombo.

Gossamer Spacecraft: Recent Trends in Design, Analysis, Experimentation, and Control

Abstract: Introduction A N emerging interest in the gossamer spacecraft community is the development and design of membrane optics that meet the stringent surface quality requirement of spaceborne telescopes. Appropriately, the development of an ultralarge, multifunctional membrane optic is being tackled head on by multiple disciplines. Strides are being made in material science, engineered actuators and sensors, and modeling techniques that can handle the unique characteristics that make gossamer structures so fascinating as well as challenging. Thorough reviews of gossamer spacecraft and related issues can be found in a few key sources. In 1995, Cassapakis and Thomas1 provided a historical perspective on the development of inflated satellite technology. Their paper covers topics such as design variables for building large, inflated craft; thoughts on new deployment and rigidization techniques; multiple applications for large, inflated craft (such as satellites, space targets, decoys, and antennae); and most importantly, lessons learned from their research and areas of research most deserving of further attention. In 2001, Jenkins et al.2 assembled a bound volume for AIAA that covers many facets of gossamer technology. The volume consists of 21 chapters devoted entirely to issues important to gossamer structures, like mechanics of membrane materials, fundamentals of membrane optics, modeling of deployment and rigidization methodologies, unique materials and their properties, and conceivable applications of ultralarge, ultralightweight craft. As a follow-up to the 2001 AIAA volume, Wada and Lou3 from the Jet Propulsion Laboratory (JPL) assembled a review of the JPL’s preflight validation tests for gossamer structures. Wada and

Mars Exploration Rover Six-Degree-of-Freedom Entry Trajectory Analysis

Abstract: The Mars Exploration Rover mission delivered the rovers Spirit and Opportunity to the surface of Mars using the same entry, descent, and landing scenario that was developed and successfully implemented by Mars Pathfinder. This investigation describes the premission trajectory analysis that was performed for the hypersonic portion of the Mars Exploration Rover entry up to parachute deployment. In this analysis, a six-degree-of-freedom trajectory simulation of the entry is performed to determine the entry characteristics of the capsules. In addition, a Monte Carlo dispersion analysis is also performed to statistically assess the robustness of the entry design to off-nominal conditions to ensure that all entry requirements are satisfied. The premission results show that the attitude at peak heating and parachute deployment are well within entry limits. In addition, the parachute deployment dynamic pressure and Mach number are also well within the design requirements.

Simulations of Solid-Propellant Rockets: Effects of Aluminum Droplet Size Distribution

Abstract: Flow modeling and simulation of solid-propellant rockets from first principles is quite challenging with several physical problems, including complex evolving geometries, turbulence, and multiphase flow with a chemically reactive disperse phase. To this end, a flexible simulation framework has been developed, in which multiphase flow computations are performed that include three-way coupling between phases (mixture-droplet-smoke), conservative coupling approach, and full heat release for the burning mechanisms. Results obtained from computations with burning aluminum droplets generating aluminum-oxide smoke are described for a generic rocket geometry. The effects of injected droplet size distribution obtained with two models are investigated and show the sensitivity of these distributions to the chamber flow dynamics, primarily at the nozzle inlet. The residence time and burning droplet diameter are verified by comparison with simple analytical predictions.

Erosion of Kapton ® H by Hyperthermal Atomic Oxygen

Abstract: Organic polymers are subject to erosion from ambient atomic oxygen in low Earth orbit. The linearity of the O-atom fluence dependence of Kapton ® H erosion and the dependence of Kapton H erosion yield on surface temperature have been investigated. Sample exposures were performed with a pulsed beam containing hyperthermal O atoms that were generated with a laser detonation source. After exposure, samples were removed from the chamber in which the exposures were performed, and postexposure analyses were performed: etch depth (profilometry) and surface topography (atomic force microscopy). A systematic set of exposures, which eroded room-temperature Kapton H from 1.4 to 25 μm, showed that the erosion yield of Kapton H is linearly dependent on O-atom fluence. This result helps validate the use of Kapton H mass loss (or erosion depth) as a linear measure of the O-atom fluence of a materials exposure. The erosion of Kapton H was strongly temperature dependent. At lower temperatures (<100◦C), the erosion yield appeared to be independent of sample temperature. However, above 100◦C, the erosion yield exhibited an Arrhenius-like temperature dependence, with an apparent activation energy of 0.31 eV. These observations suggest that O-atom-induced erosion of Kapton H proceeds through direct, nonthermal, gas-surface reactions and through reactions that depend on surface temperature.

Catalytic and Radiative Behaviors of ZrB2-SiC Ultrahigh Temperature Ceramic Composites

Abstract: Experimental investigations of radiative properties and catalytic efficiency related to atomic oxygen recombination reaction at high temperatures were conducted on sintered ZrB 2 -SiC ceramic composite designed for space applications. Total hemispherical emissivity and recombination coefficient for atomic oxygen in the range 1000-1800 K were measured. The characterization campaign was conducted using the Moyen d'Essai et de Diagnostic en Ambiance Spatiale Extreme and Moyen d'Essai Solaire d'Oxydation facilities developed at the Procedes Materiaux et Energie Solaire-Centre National de la Recherche Scientifique laboratory. Microstructural analyses prior to and after the high temperature exposure into the Moyen d'Essai Solaire d'Oxydation apparatus were also carried out using x-ray photoelectron spectroscopy, x-ray diffraction, and scanning electron microscopy. High emissivity values and low recombination coefficients were found in agreement with previous experimental studies performed on similar ceramic compounds but at lower temperatures using a different measurement technique. Samples post-analysis highlighted the oxidation-induced surface modification, and the detectable dependence of the radiative behavior upon these modifications.

Influences of Electrical Conductivity of Wall on Magnetohydrodynamic Control of Aerodynamic Heating

Abstract: Influences of the electrical conductivity of the wall of a space vehicle on the control of the aerodynamic heating in Earth-reentry flight by applying the magnetic field are numerically examined using an axisymmetric two-dimensional (r-z) thermochemical nonequilibrium magnetohydrodynamic computational fluid dynamics code. Numerical results show that when the wall of an axisymmetric blunt body is assumed to be an insulating wall, applying a dipole-type magnetic field with r and z components pushes the bow shock wave away from the blunt body and reduces the aerodynamic heating. On the other hand, when the wall is assumed to be a conducting wall, the aerodynamic heating cannot be reduced by applying the magnetic field. This is because the strong Hall electric field on the r-z plane cannot be obtained in the case of the conducting wall, so that the large electric current density in the azimuthal direction cannot be obtained and the shock wave cannot be pushed away from the blunt body.

Hyper-X Post-Flight Trajectory Reconstruction

Abstract: The formulation and development of a trajectory reconstruction tool for the NASA X-43A/Hyper-X high-speed research vehicle and its implementation for the reconstruction and analysis of flight-test data are discussed. Extended Kalman filtering techniques are employed to reconstruct the trajectory of the vehicle, based on numerical integration of inertial measurement data along with redundant measurements of the vehicle state provided by global positioning system measurements of position and velocity. The equations of motion are formulated to include the effects of several systematic error sources, the values of which may also be estimated by the filtering routines. Additionally, smoothing algorithms have been implemented in which the final value of the state (or an augmented state that includes other systematic error parameters to be estimated) and covariance are propagated back to the initial time to generate the best-estimated trajectory, based on all available data. The methods are applied to the problem of reconstructing the trajectory of the Hyper-X vehicle from flight data.

Guidance and Control Design for Hazard Avoidance and Safe Landing on Mars

Abstract: To ensure successful future Mars landing missions, the lander must be capable of detecting hazards in the nominal landing zone and maneuvering to a new and safe site. Trajectory guidance and attitude commanding are formulated for the terminal descent phase when the lander is off the parachute. The autonomous six-degree-of-freedom controls are accomplished using engines and thrusters and guided by onboard hazard-avoidance sensors. The algorithms determine the available landing zone, survey them for hazards, select the best or alternate landing site based on state estimates and available propellant, and then maneuver the lander to land safely at the selected site. Computer simulations have demonstrated the satisfactory performance of the algorithms for safe landing on Mars with assumed atmospheric environments.

Trajectory design and maneuver strategy for the MESSENGER mission to mercury

Abstract: Destined to become the first spacecraft to orbit the planet Mercury, the MESSENGER spacecraft was launched on 3 August 2004. The 6.6-year ballistic trajectory to Mercury will utilize six gravity-assist flybys of Earth (one), Venus (two), and Mercury (three). With three trajectory correction maneuvers completed by mid-December 2005, many more maneuvers will be necessary during the journey to Mercury and the subsequent 1-year duration Mercury orbit phase. The spacecraft's design and operational capability will enable real-time monitoring of every course-correction maneuver. A complex mission plan will provide multiple opportunities to obtain observational data that will help fulfill the mission's scientific objectives. Soon after entering Mercury orbit in mid-March 2011, the initial primary science orbit will have an 80-deg orbit inclination relative to Mercury's equator, 200-km periapsis altitude, 60°N subspacecraft periapsis latitude, and a 12-h orbit period. With science goals requiring infrequent orbit-phase trajectory adjustments, pairs of orbit-correction maneuvers occur at about the same time every Mercury year, or every 88 days. For the first time, the spacecraft's orbit design at Mercury accounts for the best available Mercury gravity model, small solar pressure perturbations due to changes in the solar array tilt angle, and an improved strategy for performing orbit correction maneuvers.

Entry Configurations and Performance Comparisons for the Mars Smart Lander

Abstract: The Mars Smart Lander (MSL, renamed and redefined as the Mars Science Laboratory) will provide scientists with access to previously unachievable landing sites by providing precision landing to less than 10 km of a target landing site with landing altitude capability to 2.5 km above the Mars Orbiter Laser Altimeter geoid. Precision landing is achieved by using the aerodynamic forces on the entry body to aeromaneuver through the Martian atmosphere during the entry phase of flight. The entry body is designed to provide aerodynamic lift. The direction of the aerodynamic lift vector, defined by the vehicle bank angle, is commanded by the onboard entry guidance, to converge downrange and crossrange errors by parachute deploy, while meeting the parachute deploy constraints. Several approaches and entry body configurations for providing aerodynamic lift can be considered, including axisymmetric capsule configurations with offset c.g.s using ballast or packaging, aerodynamically shaped capsule-type configurations, and alternate configurations such as mid-lift-to-drag-ratio vehicles. The design considerations, entry configurations, and entry performance of the Mars Smart Lander are described.

Micromechanical Models for Bending Behavior of Woven Composites

Abstract: Thin woven composites have been popular for space structures due to the symmetrical and balanced properties. Although in-plane properties of these materials can be calcu- lated accurately using the classical lamination theory (CLT), the corresponding bending properties lack any accuracy for one or two-ply woven laminates. Experiments on thin laminates made from woven composites disagree with the estimates of bending stifiness and strains using CLT. Such estimates can result in errors of up to 200% in the maximum bending strains or stresses, and up to 400% in the bending stifinesses. This is because CLT assumes that the flbers and the matrix are uniformly distributed in each lamina, and relies on this uniformity in the integration of the transformed laminate stifinesses over the thickness of the laminate. However, a thin laminate made from fabrics in fact con- sists of bundles of flbers that are typically much thinner than the overall thickness of the laminate; these bundles are not homogenous through the thickness. This paper presents micromechanical models for bending behavior of woven composites considering the flber bundles and the matrix and their interactions. Finite element models are developed to estimate the bending properties of plain weave composites. The results are compared to experimental data, showing very good agreement particularly for a lamina.

Aeroheating Environments for a Mars Smart Lander

Abstract: A proposed Mars Smart Lander is designed to reach the surface via lifting-body atmospheric entry (alpha = 16 deg) to within 10 km of the target site. CFD (computational fluid dynamics) predictions of the forebody aeroheating environments are given for a direct entry from a 2005 launch. The solutions were obtained using an 8-species gas in thermal and chemical nonequilibrium with a radiative-equilibrium wall temperature boundary condition. Select wind tunnel data are presented from tests at NASA Langley Research Center. Turbulence effects are included to account for both smooth body transition and turbulence due to heatshield penetrations. Natural transition is based on a momentum-thickness Reynolds number value of 200. The effects of heatshield penetrations on turbulence are estimated from wind tunnel tests of various cavity sizes and locations. Both natural transition and heatshield penetrations are predicted to cause turbulence prior to the nominal trajectory peak heating time. Laminar and turbulent CFD predictions along the trajectory are used to estimate heat rates and loads. The predicted peak turbulent heat rate of 63 W/sq cm on the heatshield leeward flank is 70% higher than the laminar peak. The maximum integrated heat load for a fully turbulent heat pulse is 38% higher than the laminar load on the heatshield nose. The predicted aeroheating environments with uncertainty factors will be used to design a thermal protection system.

Vented Tank Resupply Experiment: Flight Test Results

Abstract: This paper reports the results of the Vented Tank Resupply Experiment (VTRE) which was flown as a payload on STS 77. VTRE looks at the ability of vane propellant management devices (PMD) to separate liquid and gas in low gravity. VTRE used two clear 0.8 cubic foot tanks one spherical and one with a short barrel section and transferred Refrigerant 113 between them as well as venting it to space. Tests included retention of liquid during transfer, liquid free venting, and recovery of liquid into the PMD after thruster firing. Liquid was retained successfully at the highest flow rate tested (2.73 gpm). Liquid free vents were achieved for both tanks, although at a higher flow rate (0.1591 cfm) for the spherical tank than the other (0.0400 cfm). Recovery from a thruster firing which moved the liquid to the opposite end of the tank from the PMD was achieved in 30 seconds.

Sizing/Optimization of a Small Satellite Energy Storage and Attitude Control System

Abstract: : The recent advent of miniature single gimbal control moment gyroscopes has spawned interest in variable speed versions for combined energy storage and attitude control systems on small satellites. Although much has been studied on the theory behind such a system, little has been done In optimally sizing these actuators for small satellite applications. Therefore this paper investigates the fundamental design concepts, optimal sizing, and mission benefits for these actuators. Given a set of small satellite agility and energy storage requirements, an optimal, nonlinear programming method is applied to this problem.

Two-Dimensional Magnetohydrodynamic Simulation of a Magnetic Sail

Abstract: A magnetic sail (Magsail) is a unique deep-space propulsion system that captures the momentum of the solar wind by a large artificial magnetic field produced around a spacecraft. To clarify the momentum transfer process from the solar wind to the spacecraft, we simulated the interaction between the solar wind and the artificial magnetic field of the Magsail using the magnetohydrodynamic model. The result showed the same plasma flow and magnetic field as those of the magnetic field of the Earth; when the solar wind passes a bow shock, the solar wind is decelerated and deflected because the solar wind cannot penetrate into the magnetic field, which is called the magnetosphere around the spacecraft. The change of the solar-wind momentum resulted in a pressure distribution along the magnetopause, which is the boundary between the solar-wind plasma and the magnetosphere. The pressure on the magnetopause is then transferred to the spacecraft via the Lorentz force between the induced current along the magnetopause and the current along the coil of the spacecraft. The simulation successfully demonstrated that the change of the momentum of the solar wind is transferred to the spacecraft via the Lorentz force, and the drag coefficient of the Magsail was estimated to be 0.9 ± 0.1 when the magnetic dipole is parallel to the solar wind.

Aerodynamic Stability of Reentry Heat Shield Shapes for a Crew Exploration Vehicle

Abstract: A parametric study of the static stability of blunt-body reentry heat shield geometries applicable to a crew exploration vehicle has been performed. Performance trends are identified by varying geometric parameters that define a range of cross sections and axial shapes. Cross sections considered include oblate and prolate ellipses, rounded-edge polygons, and rounded-edge concave polygons. Axial shapes consist of the spherical segment, spherically blunted cone, and power law. Aerodynamic performance results that are based on a Newtonian surface pressure distribution have been verified against wind tunnel and flight data for the Apollo Command Module. Results are within 10% for aerodynamic coefficients, and trim angles of attack are computed within 1.2-deg. Stability and aerodynamic characteristics are observed to be more sensitive to changes in axial shape than changes in cross section. When uniform density is assumed, increased stability and performance are demonstrated at negative angles of attack for geometries with extremely blunt axial shapes and noneccentric cross sections. An unstable, oblate spherical segment at 20-deg angle of attack can produce a 56.1% increase in lift-to-drag ratio compared to a noneccentric stable spherical segment Shifting the center of gravity forward by 23.5% of its length can longitudinally stabilize this shield. The elliptical cross section, followed by the rounded-edge hexagon, and then by the rounded-edge concave hexagon rendered the most stable shapes.