Showing papers in "Journal of Spacecraft and Rockets in 2004"

Multi-Attribute Tradespace Exploration as Front End for Effective Space System Design

Abstract: The inability to approach systematically the high level of ambiguity present in the early design phases of space systems causes long, highly iterative, and costly design cycles. A process is introduced and described to capture decision maker preferences and use them to generate and evaluate a multitude of space system designs, while providing a common metric that can be easily communicated throughout the design enterprise. Communication channeled through formal utility interviews and analysis enables engineers to better understand the key drivers for the system and allows for a more thorough exploration of the design tradespace. Multi-attribute tradespace exploration with concurrent design, a process incorporating decision theory into model- and simulation-based design, has been applied to several space system projects at the Massachusetts Institute of Technology. Preliminary results indicate that this process can improve the quality of communication to resolve more quickly project ambiguity and to enable the engineer to discover better value designs for multiple stakeholders. The process is also integrated into a concurrent design environment to facilitate the transfer of knowledge of important drivers into higher fidelity design phases. Formal utility theory provides a mechanism to bridge the language barrier between experts of different backgrounds and differing needs, for example, scientists, engineers, managers, etc. Multi-attribute tradespace exploration with concurrent design couples decision makers more closely to the design and, most important, maintains their presence between formal reviews.

Shape-Based Algorithm for Automated Design of Low-Thrust, Gravity-Assist Trajectories

Abstract: Given the benefits of coupling low-thrust propulsion with gravity assists, techniques for easily identifying candidate trajectories would be extremely useful to mission designers. The computational implementation of an analytic, shape-based method for the design of low-thrust, gravity-assist trajectories is described. Two-body motion (central body and spacecraft) is assumed between the flybys, and the gravity-assists are modeled as discontinuities in velocity arising from an instantaneous turning of the spacecraft’s hyperbolic excess velocity vector with respect to the flyby body. The method is augmented by allowing coast arcs to be patched with thrust arcs on the transfers between bodies. The shape-based approach permits not only rapid, broad searches over the design space, but also provides initial estimates for use in trajectory optimization. Numerical examples computed with the shape-based method, using an exponential sinusoid shape, are presented for an Earth‐Mars‐Ceres rendezvous trajectory and an Earth‐Venus‐Earth‐Mars‐Jupiter flyby trajectory. Selected trajectories from the shape-based method are successfully used as initial estimates in an optimization program employing direct methods.

Design and Evaluation of an Acceleration Guidance Algorithm for Entry

Abstract: The design and performance evaluation of an entry guidance algorithm for future space transportation vehicles is presented. The guidance concept is to plan and track aerodynamic acceleration. This concept, on which the longitudinal entry guidance for the Space Shuttle Orbiter is based, is extended to integrated longitudinal and lateral guidance. With integrated longitudinal and lateral guidance, more extreme points in the landing footprint can be reached accurately; in particular, the cross-range capability is extended. The guidance algorithm consists of two components: a trajectory planner and a trajectory tracking law. The planner generates reference drag acceleration and heading angle profiles, along with reference state and bank angle profiles. The planner executes onboard and is capable of generating updates as the entry evolves. The tracking law, based on feedback linearization, commands the angles of bank and attack required to follow the reference drag and heading angle profiles. The planner and tracking law are described, along with additional higher level logic included in the algorithm. Extensive simulations for a set of return-from-orbit entries, including ones requiring large cross range, demonstrate that this algorithm consistently achieves the desired target conditions within allowable tolerances and satisfies all other entry constraints.

Role of Software in Spacecraft Accidents

Abstract: The first and most important step in solving any problem is understanding the problem well enough to create effective solutions. To this end, several software-related spacecraft accidents were studied to determine common systemic factors. Although the details in each accident were different, very similar factors related to flaws in the safety culture, the management and organization, and technical deficiencies were identified. These factors include complacency and discounting of software risk, diffusion of responsibility and authority, limited communication channels and poor information flow, inadequate system and software engineering (poor or missing specifications, unnecessary complexity and software functionality, software reuse without appropriate safety analysis, violation of basic safety engineering practices in the digital components), inadequate review activities, ineffective system safety engineering, flawed test and simulation environments, and inadequate human factors engineering. Each of these factors is discussed along with some recommendations on how to eliminate them in future projects.

Electromagnetic Formation Flight for Multisatellite Arrays

Abstract: The use of propellant to maintain the relative orientation of multiple spacecraft in a sparse aperture telescope such as NASA’s Terrestrial Planet Finder (TPF) poses several issues. These include fuel depletion, optical contamination, plume impingement, thermal emission, and vibration excitation. An alternative is to eliminate the need for propellant, except for orbit transfer, and replace it with electromagnetic control. Relative separation, relative attitude, and inertial rotation of the array can all be controlled by creating electromagnetic dipoles on each spacecraft, in concert with reaction wheels, and varying their strengths and orientations. Whereas this does not require the existence of any naturally occurring magnetic fields, such as the Earth’s, such fields can be exploited. Optimized designs are discussed for a generic system and a feasible design is shown to exist for a five-spacecraft, 75-m baseline TPF interferometer.

Generalized Analytical Solar Radiation Pressure Modeling Algorithm for Spacecraft of Complex Shape

Abstract: The theoretical background of solar radiation pressure modeling is presented. The attitude behavior of specific classes of spacecraft and how these attitudes make it feasible to model radiation pressure effects in the spacecraft body-fixed coordinate system are explained. The core mathematical and computational elements of the radiation pressure modeling algorithm are described. Specific components of the algorithm that enable the accurate modeling of curved surfaces and the speed optimization of the computational process are detailed. The methods and results of various forms of validation of the modeling method are given. The principal benefits of the method are that it can deal with realistic and complex spacecraft structures easily and efficiently and that it enables forms of analysis that were previously impossible. These include modeling the effects of how radiation reflected from the spacecraft surface can strike other parts of the structure causing further acceleration and simulating the effects of anomalous forces caused by departure from the spacecraft’s nominal attitude regime. Nomenclature A = surface area of a surface component, m 2 Aλ = coefficient of squared term in locus of principal section of paraboloid a = coefficient of squared term in locus of displaced paraboloid section b1‐b4 = quaternion components C = constant term in locus of principal section of paraboloid c = constant term in locus of displaced paraboloid section

Development of Advanced Metallic-Thermal-Protection System Prototype Hardware

Abstract: A new metallic thermal-protection-system concept has been designed, analyzed, and fabricated. A specific location on a slngle-stage-to-orbit reusable launch vehicle was selected to develop loads and requirements needed todesign prototype panels. The design loads include ascent and entry heating rates, pressures, acoustics, and accelerations. Additional design issues were identified and discussed. An iterative sizing procedure was used to size the thermal protection system panels for thermal and structural loads as part of an integrated wall construction that included the thermal protection system and cryogenic tank structure. The panels were sized to maintain acceptable temperatures on the underlying structure and to operate under the design structural loading. Detailed creep analyses were also performed on critical components of the panels. Four 18-in.-square metallic thermal-protection-system panels were fabricated. A lightweight, thermally compliant support system to connect the thermal protection system to the cryogenic tank structure was designed and fabricated.

Kerosene vs. Methane: A Propellant Tradeoff for Reusable Liquid Booster Stages

Abstract: Kerosene and methane are two promising candidate propellants for a future reusable booster stage. This study assesses the merits of both propellants and compares their respective performance when used in a booster stage. First of all, the principal properties of both propellants are identified. An analysis of a comparable full-flow staged combustion cycle engine for each propellant follows. The final assessment is made based on the results of a performance analysis of a launch vehicle making use of these motors in reusable fly-back boosters. The use of kerosene as propellant leads to a lower booster dry mass, making it the preferred choice if no operational benefits of methane can be identified.

New Methods for Rapid Architecture Selection and Conceptual Design

Abstract: New methods for rapid front-end development of complex systems are introduced. New tradespace exploration techniques, advances in integrated concurrent engineering, and application of risk analysis methods early in the design process allow rapid progress from poorly defined user needs to fairly detailed conceptual designs. An overview is provided of the methods. A process is described that allows thousands of system architecture alternatives to be quickly and quantitatively assessed vs user needs. The result is an understanding of the tradespace, including its key constraints and sensitivities, as well as an optimum architecture. This architecture is used to specify needs for space vehicles, which are designed using integrated concurrent engineering techniques. Research in risk and uncertainty, policy impacts, and information technology methods allows quantitative consideration of these factors, resulting in designs that are robust to uncertainties and policy impacts and potentially more versatile and flexible. Eight systems designed to date using the method are briefly reviewed. Key literature and a number of companion papers that go into depth on various aspects of the method are cited.

Multi-Objective, Multidisciplinary Design Optimization Methodology for Distributed Satellite Systems

Abstract: A multi-objective, multidisciplinary design optimization methodology for mathematically modeling the distributed satellite system conceptual design problem as an optimization problem has been developed. The tradespace for distributed satellite systems can be enormous, too large to enumerate, analyze, and compare all possible architectures. The seven-step methodology enables an efficient search of the tradespace for the best families of architectures during the conceptual design phase. Four classes of optimization techniques are investigated, Taguchi, heuristic, gradient, and univariate methods. The heuristic simulated annealing algorithm found the best distributed satellite system architectures with the greatest consistency due to its ability to escape local optima within a nonconvex tradespace. The conceptual design problem scope is then broadened by expanding from single-objective to multiobjective optimization problems, and two variant multi-objective simulated annealing algorithms are developed. Finally, several methods are explored for approximating the true global Pareto boundary with only a limited knowledge of the full design tradespace. In this manner, the methodology serves as a powerful, versatile systems engineering and architecting tool for the conceptual design of distributed satellite systems. Nomenclature Av = system availability C = capability E = system energy L = mission duration N = number of dimensions in the objective function P = candidate Pareto optimal set P ∗ = true Pareto optimal set p = state probability T = system temperature ˜ T = arithmetic mean t = time U = utility y = mission year Γ = design vector Γb = current baseline design vector Γi = initial design vector γ = design vector variable � = objective function value difference θr = angular resolution, marcsec � = life-cycle cost, $ χ = random number between 0 and 1 � = system performance � = null depth

Design and Deployment of a Satellite Constellation Using Collaborative Optimization

Abstract: A study of collaborative optimization as a systematic, multivariable, multidisciplinary method for the conceptual design of satellite constellations is presented. Collaborative optimization was selected because it is well suited to a team-oriented environment, such as often found in the constellation design process. The method provides extensive and formal exploration of the multidisciplinary design space and a scalable formulation of the problem without compromising its subsystems’ flexibility or eliminating opportunities for collaboration. The feasibility and benefits of the collaborative optimization architecture are highlighted by the successful convergence of an application notional problem to design and deploy elements of a space-based infrared system that provides early missile warning. Furthermore, this study contributes to the existing knowledge of the collaborative optimization method by verifying the feasibility of nongradient optimization algorithms as both system and subsystem optimizers within the architecture. Finally, the demonstrated convergence of this problem, which involves integer variables, also demonstrates the flexibility of the architecture for handling mixed-discrete nonlinear multidisciplinary problems.

End-to-end optimization of chemical-electric orbit-raising missions

Abstract: A simple analytic multistage model is presented for combined chemical-electric orbit-raising missions. Expressions for transportation rates and optimum electric specific impulse are derived for two-stage, three-stage, variable-efficiency, and tank-limited missions of up to 100 days duration. The optimum electric specific impulse is shown to depend strongly on the specific impulse of the chemical thruster. A low-thrust-trajectory optimization model is combined with launch-vehicle performance data to derive end-to-end optimized three-dimensional chemical-electric orbit-raising profiles to geostationary orbit. Optimized profiles are derived for the Sea Launch, Ariane 4, Atlas V, Delta IV, and Proton launch vehicles. Optimum electric orbit-raising starting orbits and payload mass benefits are calculated for each vehicle. The mass benefit is shown to be between 6.1 and 7.6 kg/day with two SPT-140 thrusters, or up to 680 kg for 90 days of electric orbit raising. The optimized profiles are combined with the analytic model to a create simple parametric performance model describing multiple launch vehicles. The model is a good tool for system level analysis of electric orbit-raising missions and is shown to match calculated performance to within 13%.

Catalytic Atom Recombination on ZrB2/SiC and HfB2/SiC Ultrahigh-Temperature Ceramic Composites

Abstract: Results are presented of an experimental investigation into the efficiency of zirconium diboride/silicon carbide and hafnium diboride/silicon carbide ultrahigh-temperature ceramic composites for catalyzing the surface recombination of dissociated oxygen and nitrogen at moderate surface temperatures. Experiments were conducted with a diffusion-tube side-arm reactor, together with laser-induced fluorescence species detection diagnostics. Experiments reveal recombination coefficients in the range between silica glasses and oxidized metals, as well as evidence of environment-induced surface modification.

Preliminary Thermal-Mechanical Sizing of a Metallic Thermal Protection System

Abstract: A process is developed to perform thermal and structural analysis and sizing and to perform sensitivity studies on the latest metallic thermal protection system developed at NASA Langley Research Center. The process defined can be used to determine appropriate materials and approximate thicknesses and is the basis for initial thermal protection system weight estimates. Metallic thermal protection systems are a key technology for reducing the cost of reusable launch vehicles, offering the combination of increased durability and competitive weights when compared to other Systems. Accurate sizing of metallic thermal protection systems requires combined thermal and structural analysis. Initial sensitivity studies were conducted using transient one-dimensional finite element thermal analysis to determine the influence of various design and analysis parameters on thermal protection system weight. The thermal analysis model was then used in combination with static deflection and failure mode analysis of the thermal protection system sandwich panel outer surface to obtain minimum weight configurations at three vehicle stations on the windward centerline of a representative reusable launch vehicle. The coupled nature of the analysis requires an iterative analysis process, which is described. Findings from the sensitivity analysis are reported, along with preliminary designs at the three vehicle stations considered.

Overview of satellite thermal analytical model

Abstract: Thermal analysis is the major engineering work throughout the entire satellite development process, with some crucial stages such as design, test, and ground operations simulation. In the formal design and verification (by test) phases, a general thermal mathematical model for the entire satellite is constructed from a combined conduction and radiation heat transfer equation with environmental heating and cooling as boundary conditions. Some representative numerical schemes with constraints used in satellite thermal analysis, as well as an introduction to the thermal model for the thermal balance test, are presented. However, the general thermal model may be too complicated or inefficient for conceptual design, test monitoring, and ground-operation simulation while developing a satellite. Therefore, simpler governing equations for pure radiation heating and cooling with exact mathematical solutions are developed to fulfill this objective at the expense of analysis accuracy. The linear approximation and exact formulation for solving this simplified problem, as well as the corresponding results shown in graphical form, are also discussed.

Bayesian Analysis of Launch Vehicle Success Rates

Abstract: In the choosing of a launch vehicle for a given mission or in the determination of insurance coverage and premiums fo rag iven launch, accurate estimates of the probability of success of the different launch vehicles provide important information. There are three general approaches to estimating the probability of launch success. The first is to use a probabilistic risk analysis, decomposing the system into its subsystems and components and estimating the probability of each of the failure modes. The second is to rely on expert judgment about the vehicle’s success rate as a whole, without a functional decomposition of the system. The third is to use statistical data about the past performance of the system to estimate the vehicle’s success rate. The focus is put on this last approach, using Bayesian probability theory to make better use of vehicle-level performance data. The procedure is demonstrated by an analysis of the success rates of most of the major families of launch vehicles currently in use in the world. A family of launch vehicles includes all variants of a particular type of vehicle from a specific manufacturer, for example, the Delta 2. For vehicles with a small number of launch attempts, the Bayesian approach provides the advantage over classic statistical approaches of yielding estimates of both the mean future frequency of success and the uncertainty about that mean.

X-38 Experimental Aerothermodynamics

Abstract: The X-38 program seeks to demonstrate an autonomously returned orbital test flight vehicle to support the development of an operational Crew Return Vehicle for the International Space Station. The test flight, anticipated in 2002, is intended to demonstrate the entire mission profile of returning Space Station crew members safely back to earth in the event of medical or mechanical emergency. Integral to the formulation of the X-38 flight data book and the design of the thermal protection system, the aerothermodynamic environment is being defined through a synergistic combination of ground based testing and computational fluid dynamics. This report provides an overview of the hypersonic aerothermodynamic wind tunnel program conducted at the NASA Langley Research Center in support of the X-38 development. Global and discrete surface heat transfer force and moment, surface streamline patterns, and shock shapes were measured on scaled models of the proposed X-38 configuration in different test gases at Mach 6, 10 and 20. The test parametrics include angle of attack from 0 to 50 degs, unit Reynolds numbers from 0.3 x 10 (exp 6) to 16 x 10 (exp 6)/ ft, rudder deflections of 0, 2, and 5 deg. and body flap deflections from 0 to 30 deg. Results from hypersonic aerodynamic screening studies that were conducted as the configuration evolved to the present shape at, presented. Heavy gas simulation tests have indicated that the primary real gas effects on X-38 aerodynamics at trim conditions are expected to favorably influence flap effectiveness. Comparisons of the experimental heating and force and moment data to prediction and the current aerodynamic data book are highlighted. The effects of discrete roughness elements on boundary layer transition were investigated at Mach 6 and the development of a transition correlation for the X-38 vehicle is described. Extrapolation of ground based heating measurements to flight radiation equilibrium wall temperatures at Mach 6 and 10 were made and generally compared to within 50 deg F of flight prediction.

Rotation Formation Flying of Three Satellites Using Tethers

Abstract: The feasibility of rotating formation flying of satellites using flexible tethers is explored. The system is composed of three satellites connected through tethers and located at the vertices of a triangle-like configuration. The satellites are modeled as point masses, and tethers are considered massless. The general formulation of the governing equations of motions of the system moving in an elliptic orbit and including different satellite masses and tether lengths is obtained through a Lagrangian approach. The open-loop tether deployment and retrieval laws have been developed. Results of numerical simulations of the nonlinear governing equations of motion of the proposed constrained system and an equilibrium analysis indicate the feasibility of achieving formation flying of three satellites in the orbital plane. Furthermore, the equilibrium analysis leads to useful design criteria in the form of inequality constraints on the system parameters. In the case when three satellites have equal masses, the critical minimum value of spin rate for system steady-spin motion in the orbital plane is found to be 0.58 times the orbital rate. Finally, the effects of various system parameters as well as the tether deployment and retrieval on the system response have been investigated.

Compressible Wall-Injection Flows in Laminar, Transitional, and Turbulent Regimes: Numerical Prediction

Abstract: Numerical simulations of compressible rocket flows are conducted in laminar, transitional, and turbulent regimes. The laminar simulation is carried out on a planar rocket flow without nozzle using the unsteady two-dimensional Navier-Stokes system. The transitional and turbulent flows are performed in three-dimensional on an extended rocket geometry with a divergent outlet using compressible large eddy simulation (LES) models. In both cases, the compressibility effect plays an important role. In the laminar case, pressure oscillation is forced at the outflow boundary. The time-averaged part of the solution is compared with the inviscid theory of compressible rocket flow of Balakrishnan et al. (Balakrishnan, G., Linan, A., and Williams F. A., "Compressibility Effects in Thin Channels with Injection," AIAA Journal, Vol.29, No. 12, 1991, pp. 2149-2154) and the oscillatory part with the acoustic layer model of Majdalani and Van Moorhem (Majdalani, J., and Van Moorhem, W. K., "Improved Time-Dependent Flowfield Solution for Solid Rocket Motors," AIAA Journal, Vol. 36, No. 2, 1998, pp. 241-248). The mean flow from the present numerical result is in better agreement with the compressible theory than the conventional Taylor's profiles (Taylor, G. I., "Fluid Flow in Regions Bounded by Porous Surfaces," Proceedings of the Royal Society of London, Series A: Mathematical and Physical Sciences, Vol. 234, 1956, pp. 456-475), as expected. The oscillatory part of the flow agrees well in the first quarter of the axial extent, near the head end. Farther downstream, the discrepancies develop rapidly between the numerical result and the acoustic-layer model. Possible causes of the difference are the effect of compressibility, which alters the local speed of sound, hence, acoustic properties, and the interference of hydrodynamic instabilities. In the transition and turbulent regimes, the dynamic LES model is applied on different resolutions. The measurements data of Traineau et al. (Traineau, J. C., Hervat, P., and Kuentzmann, P., "Cold Flow Simulation of a Two Dimensional Nozzleless Solid Rocket Motor," AIAA Paper 86-1447, June 1986) are employed for comparison purposes. The refinement study by comparison with the measurement data suggests the importance of resolving the laminar and transition region for a reliable application of LES in transitional flows. With the consideration of this aspect, LES with efficient grid size can produce resonable accuracy. Forcing hydrodynamic instabilities and a more realistic injection fluctuations model are recommended.

Atomic-Oxygen Undercutting of Protected Polymers in Low Earth Orbit

Abstract: Hydrocarbon-based polymers that are exposed to atomic oxygen in low Earth orbit are slowly oxidized into volatile gases, which results in their erosion. Atomic-oxygen protective coatings that are both durable to atomic oxygen and effective in protecting underlying polymers have been developed. However, scratches, pin window defects, polymer surface roughness, and protective coating layer configuration can result in erosion and potential failure of protected thin polymer films even though the coatings are themselves atomic-oxygen durable. Issues are presented that cause protective coatings to become ineffective in some cases yet effective in others because of the details of their specific application. Observed in-space examples of failed and successfully protected materials using identical protective thin films are discussed and analyzed. Ground laboratory atomic-oxygen testing was conducted and compared with water vapor transport analyses from a previous study of protective coatings on Kapton® (polyimide), which indicates that vapor-deposited aluminized films are not as protective as sputter-deposited silicon dioxide films because of a greater number of pin window defects. Computational modeling was conducted and indicates that atomic-oxygen atoms trapped between the front and back surface aluminized films cause accelerated undercutting damage.

Hybrid Large-Eddy / Reynolds-Averaged Navier-Stokes Simulation of Shock-Separated Flows

Abstract: An assessment of a hybrid large-eddy/Reynolds-averaged simulation (LES/RANS) procedure for high-speed, shock-separated flows is reported. A distance-dependent blending function is used to shift the turbulence closure fromMenter's k-w shear-stress-transport model near solid surfaces to a k-Δ subgrid closure away from solid surfaces and in free-shear regions. A modified recycling/rescaling procedure is used to generate time-dependent fluctuation data that are fed into the inflow plane for some calculations, with the goal being to replace the incoming boundary layer with a hybrid LES/RANS boundary layer that maintains nearly the same levels of fluctuation energy. Simulations of Mach 3 flow over a ramped-cavity configuration highlight the effects of grid refinement and choice of hybridization strategy, while simulations of Mach 3 flow over a 20-deg compression corner illustrate the effects of the choice of model constants and the inclusion of boundary-layer recycling on the mean-flow solutions.

Metallic Thermal Protection System Requirements, Environments, and Integrated Concepts

Abstract: Achieving the ultimate goal of an economically viable reusable launch vehicle will eventually require developing Federal Aviation Regulation-type performance-based requirements and certification by the Federal Aviation Administration, as is currently done for commercial transports Because the necessary requirements do not currently exist, there is no verifiable and traceable link between thermal protection system design implementation and resulting performance, safety, and cost An initial attempt has been made to outline a set of performance-based thermal protection system design requirements Critical requirements that will have a profound effect on the economic viability of a reusable launch vehicle, such as those for ground hail strike, lightning strike, bird strike, rain/rain erosion, and on-orbit dehris/micrometeoroid hypervelocity impact have been proposed In addition to design requirements, the importance of both compiling a comprehensive loads envelope and deriving time- and location-consistent loads for thermal protection system design and sizing is addressed Including ascent abort trajectories as limit-load cases and on-orbit debris/micrometeoroid hypervelocity impact as one of the discrete-source-damage cases is Imperative because of their significant impact on thermal protection system design and resulting performance, reliability, and operability General features of a suite of integrated airframe concepts is summarized, and the specific details of a metallic thermal protection system concept having design flexibility that enables weight and operability to be traded and balanced is described

Closed-Loop Missile Yaw Control via Manipulation of Forebody Flow Asymmetrics

Abstract: A high-alpha, closed-loop flow-control system for missile yaw stabilization and enhanced maneuverability was designed, developed, and successfully demonstrated in a series of open- and closed-loop experiments on a fin-less 3:1 caliber tangent ogive missile model. The active flow-control-based yaw control system consisted of eight fastresponse pressure sensors and eight deployable flow effectors arranged in concentric rings on the missile nose cone. The devices were integrated with a closed-loop controller that modulated the effectors to manipulate flow asymmetry around the missile forebody. Side forces caused by crossflow separation and forebody flow asymmetries were observed on the missile model between 40 and 60 deg alpha. Parametric studies showed that actuating flow effectors in certain configurations resulted in cancellation of large side forces associated with the natural flow asymmetry under both steady and unsteady flow conditions. Exploratory studies conducted on optimized flow effector configurations resulted in control maps with wide spectrum of positive and negative side forces for yaw modulation. Dynamic experiments successfully demonstrated the ability of the closed-loop control system to generate and maintain a range of desired yawing moments during high-alpha pitch sweeps.

Development of Protective Coatings for High-Temperature Metallic Materials

Abstract: Metallic material systems with potential for high-temperature operations are critical for many land-based and space-based systems. Advanced alloys with improved elevated temperature properties and/or reduced densities offer improved structural ef"icieney and longer service life compared to more conventional alloys. However, in extreme operating environments, these alloys require coatings for environmental protection and thermal control. We discuss some results from a program to develop ultrathin, lightweight, protective coatings applied via sol-gel techniques for some emerging high-temperature alloys. The coatings were designed to reduce oxidation, increase emittance, and reduce the catalytic efficiency for recombination of dissociated hot-gas species for the candidate materials. The alloys considered in this study include PM1000 (an oxide dispersion strengthened Ni-based alloy), 602CA Ni-based alloy, and a gamma titanium aluminide alloy. Inconel 617, a Ni-based alloy, was included as a reference. Microstructural analysis and oxidation weight gain results indicated that the coatings significantly reduced oxidation damage during extended high-temperature exposures for these alloys. In addition, one coating system was shown to improve the emittance of Inconel 617. A substantial reduction in the recombination of atomic nitrogen and oxygen at the surface of Inconel 617 substrates in a hot flowing airstream was also observed.

Optimal Conceptual Design of Two-Stage Reusable Rocket Vehicles Including Trajectory Optimization

Abstract: Many candidate concepts of reusable space transportation vehicles have been proposed around the world. This paper applies an optimization method for conceptual designs of winged fully reusable rocket two-stage-to-orbit (TSTO) vehicles in both horizontal takeoff and vertical launch styles. We first describe our methods for analyzing vehicle design. Then we integrate those methods into the optimization problem, the solution of which yields the minimized parameter, defined as the total gross weight of the first-stage booster and the second-stage orbiter. This information allows us to determine the optimal vehicle configuration and flight trajectory for a highly feasible TSTO vehicle. The optimal solutions show the necessity of lightening and miniaturizing components. Vertical launch vehicles are lighter in total gross weight than horizontal takeoff vehicles. In addition to optimizing vehicle configuration, this study also optimizes ascent and return trajectories. These optimizations enable the booster to glide back to the launch site without propellant, despite the long downrange path from the staging point of the ascent trajectory.

Unsteady Gasdynamics Effects in Pyrotechnic Actuators

Abstract: Unsteady flow effects in pyrotechnic actuators are examined quantitatively and are shown to significantly affect the performance of a pyrotechnic device when the function time or the stroke time of the device is in the order of a characteristic gasdynamics timescale of the product gases in the device. A one-dimensional "purely" gasdynamics model is developed to simulate the unsteady effects in a closed-bomb firing and in a normally open pyrotechnic valve. The model results are compared to test data and are found to agree well with the data. The results of the gasdynamics model are also compared with those of a quasi-equilibrium model, and the limitations of both models are investigated. Consequently, a nondimensional time parameter τ c is established, which can be used to approximate the level of unsteady effect on the total energy output of a given pyrotechnic device. It is concluded that the unsteady effects can be neglected with a reasonable degree of accuracy when τ c > 1.

Simultaneous Analysis of Multi-Instrument Satellite Measurements of Atmospheric Density

Abstract: When several different instruments have been used on the same satellite to measure the thermospheric density, significantly different results have been obtained. For 40 years there have been efforts to resolve these disagreements. Recent improvements in the analysis of drag coefficients and in-track winds have led us to reexamine the measurements of density made by an accelerometer, a mass spectrometer, and a density gauge deployed together on the S3-1 satellite. A simultaneous solution using data from the mass spectrometer and accelerometer significantly reduces the error caused by winds in the auroral zone during geomagnetic storms. A dramatic improvement is obtained among the three instruments by calculating the drag coefficients appropriate to the shape and altitude of the satellite instead of using the previously assumed value of 2.2 at all altitudes. A comparison of several recent studies in which 2.2 was replaced by corrected drag coefficients shows that much of the reported statistical discrepancy of 15% between density measurements and models could be removed by using appropriate drag coefficients in constructing models and in analyzing data.

Parametric Scaling Model for Nongeosynchronous Communications Satellites

Abstract: Nomenclature g = acceleration caused by gravity, m/s2 Isp = specific impulse, s Mdry = spacecraft dry mass, kg MPL = payload mass, kg Mpp = primary power mass, kg Mprop = propellant mass, kg Mwet = spacecraft wet mass, kg M0 = premaneuver spacecraft mass, kg N = number of data points Pbol = total power at beginning of life, W Peol = total power at end of life, W PPL = payload power, W Tlife = system design lifetime, years Vsat = satellite volume, m3 V = velocity increment, m/s σ = rms deviation

Tethered Planetary Capture Maneuvers

Abstract: A new concept for the application of space tethers in planetary exploration and payload transfer is presented. We propose the deployment of a payload on a spinning tether in a hyperbolic orbit to provide it with a sufficient velocity change so that it is captured in an elliptical orbit at the destination planet. This concept of using tethers for planetary capture is investigated by conducting numerical simulations of a simplified tether system. The tether mass required to prevent rupture of the tether is optimized using numerical and iterative techniques for each of the major planets in the solar system. It is demonstrated that significant mass savings can be achieved when compared to the requirements for chemical propulsion. Finally, it is shown that controlling the tether length during the maneuver can be used to correct errors in the system trajectory for both spinning and nonspinning capture cases.