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This paper examines the velocity-free attitude stabilization of an underactuated nadir-pointing rigid spacecraft in a circular orbit in the presence of time-varying disturbances. A nonlinear observer is used to estimate the angular velocity of the spacecraft to complete an output feedback control law. A sliding-mode control strategy is proposed for this stabilization problem in conjunction with a nonlinear observer. The local ultimate boundedness of the closed-loop system is rigorously proven. Numerical examples demonstrate the performance of the controller in the presence of time-varying disturbances, nonzero orbital eccentricity, inertia matrix uncertainties, and measurement noise in a sampled data implementation. In particular, even though the theoretical ultimate boundedness guarantees are local, the numerical examples demonstrate ultimate boundedness with large initial attitude errors. [ABSTRACT FROM AUTHOR]

Spacecraft relative motion planning is concerned with the design and execution of maneuvers relative to a nominal target. These types of maneuvers are frequently used in missions such as rendezvous and docking, satellite inspection, and formation flight, where exclusion zones representing spacecraft or other obstacles must be avoided. The presence of these exclusion zones leads to nonlinear and nonconvex constraints that must be satisfied. In this paper, a novel approach to spacecraft relative motion planning with obstacle avoidance and thrust constraints is developed. This approach is based on a graph search applied to a virtual net of closed (periodic) natural motion trajectories, where the natural motion trajectories represent virtual net nodes (vertices), and adjacency and connection information is determined by conditions defined in terms of safe, positively invariant lubes built around each trajectory. These conditions guarantee that transitions from one natural motion trajectory to another natural motion trajectory can be completed without constraint violations. The proposed approach improves the flexibility of a previous approach based on the use of forced equilibria and has other advantages in terms of reduced fuel consumption and passive safety. The resulting maneuvers, if planned onboard, can be executed directly or, if planned off-board, can be used to warm start trajectory optimizers to generate further improvements. [ABSTRACT FROM AUTHOR]

The design of National Aeronautics and Space Administration Orion's prelaunch navigation system is introduced, both for the first flight test, Exploration Flight Test 1, and for the first planned Orion mission, Exploration Mission 1. A detailed tradeoff of possible design decisions is discussed, and the choices made by Orion are presented. The actual performance of the navigation system during Exploration Flight Test 1 is presented together with the navigation flight-software data provided by Orion to the ground controllers in telemetry. [ABSTRACT FROM AUTHOR]

The article presents a study related to spacecraft attitude stabilization with the use of magnetorquers with separation between measurement and actuation. Topics discussed include actuation mechanisms related to spacecraft attitude stabilization; ways to spacecraft attitude control; and use of electromagnetic actuators for generation of attitude control torques on satellites flying in low earth orbits.

The article presents a study related to mass ratio of electrodynamics tether to spacecraft on deorbit stability and efficiency. Topics discussed include ways in which space activities in low Earth orbit have produced a large population of dead bodies, or debris; guidelines to remove or mitigate the space debris; and parametrical analysis of deorbit performance.

The article presents a study related to spacecraft attitude stabilization with the use of magnetorquers with separation between measurement and actuation. Topics discussed include actuation mechanisms related to spacecraft attitude stabilization; ways to spacecraft attitude control; and use of electromagnetic actuators for generation of attitude control torques on satellites flying in low earth orbits.

The article presents a study related to adaptive attitude stabilization control design for spacecraft under physical limitations. Topics discussed include role of Attitude stabilization control is an important orbital mission for spacecraft; the saturation problem for attitude control; and use of adaptive attitude tracking controller.

The high complexity of modern autonomous aerospace systems has created a need for new, powerful analysis techniques for control design and validation. This paper exploits the capabilities of continuation algorithms to derive a numerical analysis method that provides a systematic way to investigate how system properties local to the equilibrium condition around which the system is operated change throughout the operational envelope. The method is based on computing subsets of the system's equilibrium manifold with particular properties that are specified by equality constraints. The contribution of this work lies in the derivation of extended systems of equations that enable the definition of constraints on the eigenvalues and eigenvectors of the linearized dynamics, so that sets of equilibrium points with particular characteristics of the local dynamical behavior can be computed. The capabilities of the method are demonstrated through application to a fully nonlinear aircraft model with six degrees of freedom. [ABSTRACT FROM AUTHOR]

Reliability-based robust design optimization of modern small satellites has recently been receiving much attention. The role of multidisciplinary design optimization considering system uncertainty is increasingly being recognized in improving satellite performance, safety, and reliability. However, high-dimensional uncertainty hampers the efficiency and accuracy of reliability-based robust design optimization. In this study, a novel multidisciplinary optimization methodology suitable for high-dimensional uncertainty is established and verified. The in-loop uncertainty quantification employing active subspace identification makes the uncertainty propagation and management much easier. Based on the discovered active subspace and a system decoupling strategy, a multiobjective alliance algorithm is presented for the complex constrained design optimization. Typical uncertain factors in small satellites are characterized and interdisciplinary relations are analyzed. A standard test example of speed reducer design first proves the methodology, and then an Earth observation satellite fully illustrates the efficacy. Through 23-dimensional uncertainty represented by one dimension, the multiobjective alliance algorithm gives well-distributed Pareto-optimal (PO) solutions, providing a fruitful reference for decision makers compromising different design targets, like satellite mass, total cost, and ground resolution. The proposed approach is proved to be fairly accurate and adaptable, which can be widely applied to uncertainty-based multidisciplinary design optimization (UMDO) of satellite systems. [ABSTRACT FROM AUTHOR]

The article presents a study which examined the use of an optimization model in the hypothetical operation of future space station for China, based on crew size, duration of the manned space station, upload of resources and number of manned and cargo vehicles. Topics discussed include the definition of the total crew onboard duration and mass of materials and instruments, the optimization model for the design of a space station operational scenario, and the cost of cargo vehicle.

Ceramic tiles provide lightweight, reusable insulation for spacecraft thermal protection systems but are vulnerable to orbital debris impact damage. Because a wide range of impact conditions that may be experienced in orbit cannot be duplicated in the laboratory, design for orbital debris impact effects must include analytical or numerical modeling. Recent research has developed and validated a new computational approach to the simulation of orbital debris impact on porous ceramic thermal protection systems. Simulations using the validated formulation suggest scaling rules for application in spacecraft design. [ABSTRACT FROM AUTHOR]

Well before Curiosity's successful landing at Gale Crater on the night of 5 August 2012, the entry, descent, and landing team had to contend with a number of development challenges that threatened the system architecture, the spacecraft's performance, and its safety. Given the ambitiousness of the landing system, perhaps it comes as no surprise that these challenges existed in all phases of system development, spanning design, testing, and operations and across all portions of flight from the top of the atmosphere through touchdown. These challenges, which included hardware, software, and flight dynamics issues, required a variety of different responses and wide range of expertise within the team. The ways the team answered the challenges were the difference between success and failure in both reaching the launch pad and reaching the surface of Mars. [ABSTRACT FROM AUTHOR]

Time-accurate dynamic simulation of objects and spacecraft in rarefied and continuum-rarefied transition flows during atmospheric entry is computationally challenging. The traditional ballistic approach based on averaged empirical coefficients cannot account for nonlinear effects from high-frequency attitude variation. The direct simulation Monte Carlo method is shown to provide the capability to model these unsteady effects. Time-accurate simulations of two distinct geometries in low thermospheric orbits and atmospheric entry are performed. Characteristic qualities of motion, such as stability, equilibria, and other limiting behavior, are identified. This approach illustrates the potential for accurate prediction of the unsteady orbital and entry dynamics for improved spacecraft design, atmospheric science, and orbital decay prediction as well as a number of new challenges for future work to stimulate the development of a new subfield of spacecraft dynamics. [ABSTRACT FROM AUTHOR]

This paper investigates the velocity-measurement-free feedback control problem associated with attitude tracking for a rigid spacecraft in the presence of both actuator failures and actuator saturation. The decreased reaction torque fault and the increased bias torque fault in the reaction wheel are handled. By using only the measurable attitude, a terminal sliding-mode-based observer is developed to reconstruct all the states of the attitude system in finite time. With the reconstructed information, a novel attitude tracking controller is developed. A Lyapunov analysis shows that the desired attitude trajectories are followed even in the presence of external disturbances. The key features of the proposed control approach are that it is independent from the knowledge of actuator faults and fault-tolerant control is achieved without the need of angular velocity. The attitude tracking performance using the proposed strategy is evaluated through a numerical example. [ABSTRACT FROM AUTHOR]

A method is presented to optimize the orientation of directional sensors and instruments in a vehicle body-fixed frame. Directional dependence is included by creating a uniformly distributed set of directions in the body-fixed frame and formulating the objective as a function of these directions. The method is demonstrated by application to photodiodes for sun sensing, for which the covariance of the sun vector estimate is derived as a function of the photodiode configuration. The measured sun vector angular accuracy is then minimized as a function of the configuration, which enables the most accurate sun sensing with the given hardware. This technique maximizes subsystem performance and provides a design method to replace traditional, iterative design approaches to sensor placement. [ABSTRACT FROM AUTHOR]

A novel saturated proportional-derivative control law incorporated with closed-loop control allocation is proposed for spacecraft attitude stabilization in this paper. More specifically, a saturated proportional-derivative-based baseline nonlinear controller is designed to guarantee the globally asymptotic stability under control input/signal constraints. Then, a closed-loop constrained optimal control allocation scheme is employed to distribute the moments synthesized by the baseline controller over the redundant actuators in the presence of constraints due to actuator amplitude and rate constraints. The optimal control solution is to be found by penalizing the control allocation errors and rates using optimal quadratic programming algorithms. A significant feature of this work is that the asymptotic stability with the closed-loop control allocation is guaranteed theoretically. Simulation results are presented to illustrate the performance of the proposed scheme. [ABSTRACT FROM AUTHOR]

This paper develops the analytical solution to the nondrifting transfer problem for rendezvous in elliptical orbits. The transfer algorithm generates a nondrifting, passively safe trajectory between two spacecraft slates by finding the appropriate maneuvers and transfer time. The nondrifting solution is derived first for Cartesian state vectors, which is useful for transferring to a terminal approach point that is defined in terms of such a Cartesian state vector. Next, the transfer problem is solved in terms of differential orbital elements. This second solution lends itself well to perform close-range rendezvous in elliptical orbits. [ABSTRACT FROM AUTHOR]

Motivated by aerospace applications, this paper presents a methodology to use second-order cone programming to solve nonconvex optimal control problems. The nonconvexity arises from the presence of concave state inequality constraints and nonlinear terminal equality constraints. The development relies on a solution paradigm, in which the concave inequality constraints are approximated by successive linearization. Analysis is performed to establish the guaranteed satisfaction of the original inequality constraints, the existence of the successive solutions, and the equivalence of the solution of the original problem to the converged successive solution. These results lead to a rigorous proof of the convergence of the successive solutions under appropriate conditions as well as non-conservativeness of the converged solution. The nonlinear equality constraints are treated in a two-step procedure in which the constraints are first approximated by first-order expansions, then compensated by second-order corrections in each of the successive problems for enhanced convergence robustness. Applications in highly constrained spacecraft rendezvous and proximity operations, finite-thrust orbital transfers, and optimal launch ascent are provided to demonstrate the effectiveness of the methodology. [ABSTRACT FROM AUTHOR]

The feasibility of flying a crewed lifting body, such as the HL-20, during entry from low Earth orbit without the steady-state body flap deflections required for angle-of-attack control was evaluated. This entry strategy mitigates the severity of the aerothermal environment on the vehicle's body flaps and reserves control power for transient steering maneuvers. A real-time numeric predictor-corrector entry guidance algorithm was developed to accommodate the range of vehicle trim angle-of-attack profiles possible in the absence of angle-of-attack control. Results show that it is feasible to steer the vehicle from low Earth orbit to a desired target with real-time guidance while satisfying a reasonable suite of trajectory constraints, including limits on peak heat rate, peak sensed deceleration, and integrated heat load. Uncertainty analyses confirm this result and show that the vehicle maintains significant performance robustness to expected day-of-flight uncertainties. Additionally, parametric scans over mission design parameters of interest indicate that a high level of flexibility is available for the low-Earth-orbit return mission. Together, these results indicate that the proposed entry strategy is feasible: Crewed lifting bodies may be effectively flown without steady-state body flap deflections. [ABSTRACT FROM AUTHOR]

The article presents information on "Human Exploration of Mars Design Reference Architecture 5.0," NASA's current human Mars exploration design. It references a presidential report on "flexible path to Mars" which is considered as an alternative to DRA 5.0, studied and refined over several iterations, and is considered a shared human and robotic Martian exploration campaign. The advantages and disadvantages of the mission design as well as alternative Mars exploration campaign are discussed.

This paper provides a detailed mission analysis and systems design of a pole-sitter mission. It considers a spacecraft that is continuously above either the North or South Pole and, as such, can provide real-time, continuous, and hemispherical coverage of the polar regions. Two different propulsion strategies are proposed, which result in a near- term pole-sitter mission using solar-electric propulsion and a far-term pole-sitter mission, in which the electric thruster is hybridized with a solar sail. For both propulsion strategies, minimum propellant pole-sitter orbits are designed. Optimal transfers from Earth to the pole sitter are designed, assuming Soyuz and Ariane 5 launch options, and a controller is shown to be able to maintain the trajectory under unexpected conditions, such as injection errors. A detailed mass budget analysis allows for a tradeoff between mission lifetime and payload mass capacity, and candidate payloads for a range of applications are investigated. This results in a payload of about 100 kg that can operate for approximately four years with the solar-electric spacecraft, whereas the hybrid-ropulsion technology enables extending the missions up to seven years. Transfers between north and south pole-sitter orbits are also considered to observe either pole when illuminated by the sun. [ABSTRACT FROM AUTHOR]

This paper develops a problem formulation that can be applied to the integrated space-station system design. The integrated space-station optimization formulation combines both a space habitat and its supporting logistics design into one single optimization problem. In this way, a space-based habitat and its logistics resupply system can be designed concurrently, and effective trades can be conducted between them. Numerical results show the benefits of the proposed problem formulation, comparing it with conventional design strategies. Data from the 2012 International Space Station reference configuration are used to develop four cases, each emphasizing a different mix of scientific experiments in orbit. [ABSTRACT FROM AUTHOR]

The modal assurance criterion and normalized cross-orthogonality check are widely used to assess the correlation between the experimentally determined dynamic characteristics and the finite element model predictions. In this paper, the effectiveness of these criteria on the base excitation responses of three spacecraft models is carried out. The dynamic characteristics obtained from a nominal finite element model are considered as experimental or true characteristics, and those obtained from a model produced by introducing errors in the nominal model are considered as analytically predicted characteristics. It is observed that these criteria are not suitable, particularly when the model is used to predict forced response characteristics such as the force transmitted to the base, peak acceleration response, and dynamic displacement in the spacecraft. Thus, a qualitative indicator named as base-force assurance criterion is defined by comparing the experimentally determined dynamic force at the base and the finite element predicted force such that the criterion can state the possible error in the peak acceleration and the dynamic displacement under the base excitation. The method is applied to assess the performance of three spacecraft structures, and the results show that new criterion can better correlate with the acceleration and the dynamic displacement error than the conventional criteria. [ABSTRACT FROM AUTHOR]

This paper addresses the preliminary robust design of a small/medium-scale reentry unmanned space vehicle. A hybrid optimization technique is proposed that couples an evolutionary multi-objective algorithm with a direct-transcription method for optimal control problems. Uncertainties on the aerodynamic forces and vehicle mass are integrated in the design process, and the hybrid algorithm searches for geometries that 1) minimize the mean value of the maximum heat flux, 2) maximize the mean value of the maximum achievable distance, and 3) minimize the variance of the maximum heat flux. The evolutionary part handles the system-design parameters of the vehicle and the uncertain functions, while the direct-transcription method generates optimal control profiles for the reentry trajectory of each individual of the population. During the optimization process, artificial neural networks are used to approximate the aerodynamic forces required by the direct-transcription method. The artificial neural networks are trained and updated by means of a multifidelity, evolution-control approach. [ABSTRACT FROM AUTHOR]

A systematic methodology is presented for defining a minimum functionality baseline configuration of a human spacecraft. To estimate a lower bound for the spacecraft mass, a set of essential functions is coupled to single-string subsystems with zero fault tolerance. This minimum functionality baseline is defined to meet the physical requirements needed to transport the crew to the target destination and to ensure that their physiological needs are met, but without margin, dispersions, redundancy, or factor of safety. This constitutes a set of nonnegotiable requirements based on fundamental parameters derived from physics and physiology. By definition, this represents a technically feasible solution, but it results in the highest-risk design. Mass additions beyond the minimum functional configuration are allocated to increased safety or operability through the addition of component redundancy, fault tolerance, factor of safety, additional mission functionality, and improved human-system interfaces. This proposed methodology was used to analyze a range of lunar ascent stage spacecraft configurations, and a process was developed to allow systematic estimation of mass for the spacecraft subsystems. The modeled results are verified by comparing them with actual subsystem mass of the Apollo lunar module ascent stage. [ABSTRACT FROM AUTHOR]

The article offers information on a study which explores the dynamics of the gyro-gravitational stabilized spacecraft in the mode of the deployment of the flexible pantograph structure. Dynamic behavior of the design at various values of parameters of spacecraft were analyzed with flexible pantograph structures and the laws of deployment.

The article discusses a study that developed a spacecraft model to determine the applicability of the trade space exploration process to the problem of designing for the space environment. It is said that the process will help designers understand the complexity of spacecraft systems because it allows them to change and update a spacecraft model during the structural optimization process. Noted are the effects of the space environment on the survivability of a spacecraft.

A tumbleweed rover is a spherical wind-driven rover designed to explore places of geological interest on the Martian surface. Dynamic models developed for an individual rover are used to create numerical simulations for a rover traversing through flat terrain, a channel, and a crater. The simulations show that the rover's motion is dependent on the terrain type and initial and atmospheric conditions. The results confirm that the wind force both pushes and hinders the rover's motion while sliding, rolling, and bouncing. The rover periodically transitions between these modes of movement when contact is initiated against sloped portions of terrain. Combinations of rolling and bouncing may be a more effective means of transport for a rover traveling through a channel when compared to rolling alone. The aerodynamic effects of drag and the Magnus force are contributing factors to the possible capture of the rover by a crater. [ABSTRACT FROM AUTHOR]

A program of research, development, test, and evaluation is planned for the development of spacecraft handling qualities guidelines. In this first experiment, the effects of reaction control system design characteristics and rotational control laws were evaluated during simulated proximity operations and docking. Also, the influence of piloting demands resulting from varying closure rates was assessed. The pilot-in-the-loop simulation results showed that significantly different spacecraft handling qualities result from the design of the reaction control system. In particular, cross coupling between translational and rotational motions significantly affected handling qualities as reflected by Cooper-Harper pilot ratings and pilot workload, as reflected by Task Load Index ratings. This influence is masked (but only slightly) by the rotational control system mode. Although rotational control augmentation using rate command/attitude hold can reduce the workload (principally, physical workload) created by cross coupling, the handling qualities are not significantly improved. The attitude and rate deadbands of the rate command/attitude hold introduced significant mental workload and control compensation to evaluate when deadband firings would occur, assess their impact on docking performance, and apply control inputs to mitigate that impact. [ABSTRACT FROM AUTHOR]

An evaluation of the Stardust spacecraft forebody heat shield design was conducted to address uncertainties in original development and qualification efforts. Results are reported for additional arc jet tests and analyses done on the ablative and thermal performance of the low density ablative material used in the Stardust design for conditions simulating nominal peak convective heating conditions expected for Stardust sample return capsule Earth entry. Test data were used to adjust iteratively the thermophysical properties in an ablative response computer code to match experimental in-depth temperatures. Ablative recession rates and maximum internal temperatures were satisfactorily predicted by the computer code using derived properties for nominal peak entry heating rates and for rates 37% greater than nominal. Experimental and computer model maximum internal material temperatures were in agreement over a range of conditions and are considered to validate the Stardust heat shield design although measured in-depth temperature data show consistent temperature rise deviations that are not accurately modeled by the computer code. Predicted Stardust heat shield performance based on these results indicates that maximum attachment bond line temperatures will be considerably less than the maximum allowable of 250°C, and total recession will be a small fraction of the as-designed thickness. [ABSTRACT FROM AUTHOR]

Schedule slippage plagues the space industry and is antinomic with the recent emphasis on space responsiveness. The Government Accountability Office has repeatedly noted the difficulties encountered by the U.S. Department of Defense in keeping its acquisition of space systems on schedule, and they identified the low Technology Readiness Level of the system/payload under development as a principal culprit driving schedule risk and slippage. In this paper, based on data from past space programs, we analyze the relationship between technology uncertainty and schedule risk in the acquisition of space systems and propose an analytical framework to identify appropriate schedule margins for mitigating the risk of schedule slippage. We also introduce the Technology Readiness Level schedule-risk curves to help program managers make risk-informed decisions regarding the appropriate schedule margins for a given program or the appropriate Technology Readiness Level to consider if the program's schedule were to be exogenously and rigidly constrained. Based on our findings, we recommend that the industry adopt and develop schedule-risk curves (instead of single-schedule point estimates), that these schedule-risk curves be made available to policy- and decision-makers in acquisition programs, and that adequate schedule margins be defined according to an agreed-upon and acceptable schedule-risk level. [ABSTRACT FROM AUTHOR]

Boundary-layer transition continues to be a critical factor in hypersonic fight vehicle design. Measurements of transition during hypersonic flight testing provide valuable data for the development and verification of transition-prediction techniques. A summary of transition-measurement techniques used on vehicles flown by Sandia National Laboratories is presented, including sample flight data to illustrate the type of transition indication obtained from each measurement technique. The Slender Hypervelocity Aerothermodynamic Research Probe, a ballistic vehicle flown by Sandia for NASA, is used as a case study to illustrate how transition is determined for a flight vehicle and to illustrate some of the difficulties associated with these types of measurements. [ABSTRACT FROM AUTHOR]

High reliability is a primary design goal in commercial communication satellite systems because they require large capital investment and are inaccessible after launch. Spacecraft system reliability is typically computed using standard parallel-series combination techniques based upon component and subsystem failure rates provided by suppliers. Component failure rates are empirically determined and, as such, are nondeterministic parameters. Treating these failure rates as uncertain parameters in spacecraft design may avoid unnecessary high redundancy. However, probabilistic methods in reliability evaluation are often ignored because handling uncertain parameters associated with discrete or categorical design variables, such as technology choices and redundancy levels, requires computationally expensive sampling techniques. The computational cost of optimization approaches becomes prohibitive when considering discrete technology and redundancy choices as variables. This work presents a genetic algorithm with Monte Carlo sampling for probabilistic reliability-based design optimization of satellite systems. In this approach, confidence-level constraints ensure that system reliability requirements are met with high probability. The genetic algorithm-Monte Carlo sampling approach is compared to a deterministic margin-based approach that enforces margins or safety factors on the reliability of individual components. The comparison shows that the genetic algorithm-Monte Carlo sampling approach produces satellite designs that have low launch mass (a surrogate for cost) while achieving reliability requirements at specified high confidence levels, while the genetic algorithm--deterministic margin-based approach produces heavy satellite designs with excessive redundancy. Based on this work, extensions of a genetic algorithm-based approach for discrete optimization under uncertainty that may require less computational effort appear possible. [ABSTRACT FROM AUTHOR]

Focuses on the development of a low-diffusion efficient upwind scheme for compressible flow governing equations. Importance of the scheme for aircraft and spacecraft design; Modification of the original E-CUSP scheme to remove the temperature oscillation with numerical dissipation; Comparison between the modified scheme and Roe scheme.