Venus and Beyond Using the Ariane ASAP Launch Capability
A general trajectory scenario has been designed which can transfer a spacecraft from a pre-specified, highly eccentric, near equatorial orbit to a desired escape direction.
Venus↗
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Search indexed NASA NTRS and DOE OSTI research on propulsion, heat transfer, battery materials and energy systems. Follow report and document links to the original sources.
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Anisotropic angular scattering models of elastic electron-neutral collisions for Monte Carlo plasma simulations
Many laboratory and industrial plasma applications require accurate modeling techniques to understand the interplay between microscopic and macroscopic processes. A prime example of this interplay is how particle and Monte Carlo (MC) simulation codes describe angular scattering of electrons following elastic scattering events. The forward peaked nature of high energy electron elastic scattering is relatively trivial to accurately describe in plasma simulations. However, for lower energy collisions, which produce near isotropic or backward peaked differential cross sections, there is not a strong consensus among the plasma modeling community on how to best describe these angular scattering trends. Here, in this study, we propose a systematic method to approximate the aforementioned non-trivial angular scattering behavior with a formula that can be readily implemented in particle-in-cell (PIC) and/or MC plasma simulation codes. The present approach is specifically applied to fusion relevant atomic hydrogen and helium, as well as for molecular hydrogen, and results are also applicable to the atomic isotopes and homonuclear molecular isotopologues of these species. Comparisons between the present angular distribution function and benchmark scattering data were used to validate the proposed models. In addition, two-term Boltzmann calculations and PIC direct simulation MC simulations revealed that the proposed angular distribution function is accurate, agreeing very well with benchmark convergent close-coupling scattering calculations, and electron transport measurements. These studies confirmed that the present angular distribution function model can be utilized without the need of renormalization to the momentum transfer cross section (as opposed to using the elastic scattering integrated cross section), which has been suggested by several studies in order to correct for deficient angular scattering models, and to agree with transport measurements. Hence, the present anisotropic angular scattering model can be utilized to accurately model the momentum transfer as well as the electron trajectories of elastic collisions.
70 PLASMA PHYSICS AND FUSION TECHNOLOGY↗
Conditions for optimal one-impulse transfer.
Optimal one- and two-impulse orbital transfer maneuvers, considering shallowly intersecting coplanar orbits
OPTIMUM THRUST PROGRAMMING↗
Optimum deorbit positioning for minimum impulse reentry.
Optimum deorbit positioning on planetocentric elliptic orbit for single impulse reentry treated by two coupled quartic polynomials
REENTRY PHYSICS↗
TROJID: A portable software package for upper-stage trajectory optimization
Performance optimization for upper-stage exoatmospheric vehicles often is performed within the framework of a full capability trajectory simulation package requiring either a large mainframe computer or powerful work-station. Since these software packages tend to include capabilities providing for high-fidelity boost and reentry simulations, the programs usually are quite large and not very portable. The program TROJID is an attempt to provide an environment for the optimization of upper-stage trajectories within a small package capable of being run on a standard desktop microcomputer. Utilizing a state-of-the-art nonlinear programming algorithm and a trajectory simulator implementing impulsive burns and an analytic coast phase propagator, TROJID is capable of producing trajectories for optimal multi-burn upper-stage orbit transfers. The package has been designed to allow full generality in definition of both the trajectory simulator and the parameter optimization problem.
Hammes, Steven M.↗
Aerothermal Analysis of the Project Fire II Afterbody Flow
Computational fluid dynamics (CFD) is used to simulate the wake flow and afterbody heating of the Project Fire II ballistic reentry to Earth at 11.4 km/sec. Laminar results are obtained over a portion of the trajectory between the initial heat pulse and peak afterbody heating. Although non-catalytic forebody convective heating results are in excellent agreement with previous computations, initial predictions of afterbody heating were about a factor of two below the experimental values. Further analysis suggests that significant catalysis may be occurring on the afterbody heat shield. Computations including finite-rate catalysis on the afterbody surface are in good agreement with the data over the early portion of the trajectory, but are conservative near the peak afterbody heating point, especially on the rear portion of the conical frustum. Further analysis of the flight data from Fire II shows that peak afterbody heating occurs before peak forebody heating, a result that contradicts computations and flight data from other entry vehicles. This result suggests that another mechanism, possibly pyrolysis, may be occurring during the later portion of the trajectory, resulting in less total heat transfer than the current predictions.
Wright, Michael J.↗
Q-Law for Rapid Assessment of Low Thrust Cislunar Trajectories Via Automatic Differentiation
Q-Law is a Lyapunov-based control law used to determine optimal controls for a low thrust trajectory. One major issue with its use is the difficult derivatives re-quired for calculating optimal controls at a given time. In this paper, an implementation of Q-Law with automatic differentiation via a Python package called JAX is applied. With automatic differentiation, the difficult derivatives for Q-Law’s optimal controls are calculated with ease, and derivatives of final states with respect to Q-Law’s weights are found enabling gradient-based optimization of Q-Law for the first time. Different search and optimization methods for finding optimal weights are then compared using the LEO to GEO problem, and it was found that gradient-free methods like design of experiments and genetic algorithm produced the best results, but they took the longest time to get a solution, while the gradient-based method found a locally optimal result in a much faster time. Overall, the run time for a single propagation is manageable and well-suited for a mission designer to use as an initial guess generator for trajectory optimization, or for simple orbit transfer analysis.
Nathan Steffen↗
Integrated Vehicle and Trajectory Design of Small Spacecraft with Electric Propulsion for Earth and Interplanetary Missions
This paper investigates the feasibility of Earth-transfer and interplanetary mission architectures for miniaturized spacecraft using emerging small solar electric propulsion technologies. Emerging small SEP thrusters offer significant advantages relative to existing technologies and will enable U-class systems to perform trajectory maneuvers with significant Delta V requirements. The approach in this paper is unique because it integrates trajectory design with vehicle sizing and accounts for the system and operational constraints of small U-class missions. The modeling framework includes integrated propulsion, orbit, energy, and external environment dynamics and systems-level power, energy, mass, and volume constraints. The trajectory simulation environment models orbit boosts in Earth orbit and flyby and capture trajectories to interplanetary destinations. A family of small spacecraft mission architectures are studied, including altitude and inclination transfers in Earth orbit and trajectories that escape Earth orbit and travel to interplanetary destinations such as Mercury, Venus, and Mars. Results are presented visually to show the trade-offs between competing performance objectives such as maximizing available mass and volume for payloads and minimizing transfer time. The results demonstrate the feasibility of using small spacecraft to perform significant Earth and interplanetary orbit transfers in less than one year with reasonable U-class mass, power, volume, and mission durations.
Spangelo, Sara↗
Electric Sail Design Sensitivities
Electric Sails (E-Sails) are a promising propulsion technology that seek to enable high characteristic propellentless acceleration for spacecraft to reach distant and/or difficult to reach orbits such as rapid transit to heliopause. The E-Sail system exchanges momentum by using positively charged electrostatic tethers to repel solar wind photons to push it through space. This concept was first theorized by Pekka Janhunen in 2004 with further developments occurring including a NASA Innovative Advanced Concepts Phase 1 and Phase 2 hosted out of NASA’s Marshall Space Flight Center (MSFC). These developments led to further maturation of the overall system, and this paper is designed to help identify different design sensitivities of an integrated Electric Sail system. To approach this, a small team at MSFC took an in-house developed three degrees of freedom (3DoF) simulation tool and a trajectory modelling tool to look at different design parameters and to determine a potential ideal E-Sail configuration. An E-Sail system has different design architectures including a barbell design, hub and spoke design, and a potential hybrid solution. The barbell design features two equally massed satellites that spin around a central point in the tether system. The hub and spoke design features a large central spacecraft with small end spacecraft to aid in formation control of the overall E-Sail system. Leveraging elements from both the hub and spoke, and barbell design, a hybrid option exists where one could have a larger central mass and one or two tethers extended to a smaller end mass [2]. In summary, the hub and spoke design is the ideal configuration for E-Sail with tethers spanning kilometers to achieve the designed design characteristic acceleration of at least one mm/s2 with this architecture being the focus of this study. Different key design parameters were varied as part of this study. These parameters include the total number of tethers, the length of the tethers, spin rate, relative spacecraft mass, tether voltage, inertia per tether, and impact of slew rate changes. These results started from an internal MSFC technology demonstration mission design and then the parameters were varied with engineering judgement to ensure that the results were consistent with expected results. These were varied in the MSFC 3DoF tool and compared to baseline results to generate a candidate ideal E-Sail design system. These results were then used to help inform the mission analysis design for the system. The 3DoF trajectory optimization was performed in the Astrodynamics and Space Science Enabling Toolbox (ASSET) and utilized collocation optimization of control vector pointing with the objective function being transfer duration. A representative model of the E-Sail force was derived as a function of the Sun-pointing vector direction. This model was initialized with a zero Sun Incidence Angle (SIA) trajectory for the initial guess generation, then with a portion of the mission optimized SIA time history for fast outbound transfers. After the trajectory is hyperbolic relative to a heliocentric frame, the SIA is maintained at 0 degrees except for any needed trajectory correction maneuvers. This trajectory optimization looked at different characteristic acceleration values to analyze the impact it had on mission performance. As a result of this study, the team can best inform mission designers and technology development efforts to mature the E-Sail system. This study will allow mission designers to have defined rules-of-thumb to design an E-Sail system to meet desired mission needs as well as providing several sample mission profiles. These design drivers will inform maturation efforts on which design requirements to consider for an integrated E-Sail design.
E-Sail↗
Optimal orbital transfers and prime rib curves for mission design
Battin's work on optimal one impulse transfers from circular orbits is extended. Optimal one impulse trajectories from orbits of arbitrary eccentricity through a specified point are determined. This work is then further extended to optimal two impulse transfers between massive planets. Curves derived using this technique are presented showing flyby and rendezvous Delta V requirements for all combinations of launch and arrival true anomalies. Similar curves for times of flight corresponding to the trajectories with minimum total Delta V are derived using Lambert's methods. Possible further extension of these techniques are outlined. The use of these curves for mission design is presented in a companion paper by Hulkower and Ross (1971).
Ross, D. J.↗
The effect of interplanetary trajectory options on a manned Mars aerobrake configuration
Manned Mars missions originating in low Earth orbit (LEO) in the time frame 2010 to 2025 were analyzed to identify preferred mission opportunities and their associated vehicle and trajectory characteristics. Interplanetary and Mars atmospheric trajectory options were examined under the constraints of an initial manned exploration scenario. Two chemically propelled vehicle options were considered: (1) an all propulsive configuration, and (2) a configuration which employs aerobraking at Earth and Mars with low lift/drag (L/D) shapes. Both the interplanetary trajectory options as well as the Mars atmospheric passage are addressed to provide a coupled trajectory simulation. Direct and Venus swingby interplanetary transfers with a 60 day Mars stopover are considered. The range and variation in both Earth and Mars entry velocity are also defined. Two promising mission strategies emerged from the study: (1) a 1.0 to 2.0 year Venus swingby mission, and (2) a 2.0 to 2.5 year direct mission. Through careful trajectory selection, 11 mission opportunities are identified in which the Mars entry velocity is between 6 and 10 km/sec and Earth entry velocity ranges from 11.5 to 12.5 km/sec. Simulation of the Earth return aerobraking maneuver is not performed. It is shown that a low L/D configuration is not feasible for Mars aerobraking without substantial improvements in the interplanetary navigation system. However, even with an advanced navigation system, entry corridor and aerothermal requirements restrict the number of potential mission opportunities. It is also shown that for a large blunt Mars aerobrake configuration, the effects of radiative heating can be significant at entry velocities as low as 6.2 km/sec and will grow to dominate the aerothermal environment at entry velocities above 8.5 km/sec. Despite the additional system complexity associated with an aerobraking vehicle, the use of aerobraking was shown to significantly lower the required initial LEO weight. In comparison with an all propulsive mission, savings between 19 and 59 percent were obtained depending upon launch date.
Braun, Robert D.↗
Missed Thrust Analysis and Design for Low Thrust Cislunar Transfers
This paper details the analysis of the missed thrust problem for low thrust cislunar transfers. Missed thrust analysis is completed for a reference NRHO to DRO transfer that is designed to maximize the final mass without consideration of robustness to unexpected loss of thrust. A new missed thrust design method is presented to include the robustness of the transfer as part of the optimization problem by including a minimal number of branching trajectories tied to the start of thrust arcs in the reference transfer and optimizing them all simultaneously. Results from variations of this approach are presented for and compared to the reference transfer designed without consideration for missed thrust. The results show that this new method can reduce the additional propellant required for recovery from an unexpected 7-day outage by up to 90% without significant increase to the propellant required to complete the reference transfer.
Grebow, Daniel J.↗
Spacecraft Charging Issues for Launch Vehicles
Spacecraft charging is well known threat to successful long term spacecraft operations and instrument reliability in orbits that spend significant time in hot electron environments. In recent years, spacecraft charging has increasingly been recognized as a potentially significant engineering issue for launch vehicles used to deploy spacecraft using (a) low Earth orbit (LEO), high inclination flight trajectories that pass through the auroral zone, (b) geostationary transfer orbits that require exposures to the hot electron environments in the Earths outer radiation belts, and (c) LEO escape trajectories using multiple phasing orbits through the Earths radiation belts while raising apogee towards a final Earth escape geometry. Charging becomes an issue when significant areas of exposed insulating materials or ungrounded conductors are used in the launch vehicle design or the payload is designed for use in a benign charging region beyond the Earths magnetosphere but must survive passage through the strong charging regimes of the Earths radiation belts. This presentation will first outline the charging risks encountered on typical launch trajectories used to deploy spacecraft into Earth orbit and Earth escape trajectories. We then describe the process used by NASAs Launch Services Program to evaluate when surface and internal charging is a potential risk to a NASA mission. Finally, we describe the options for mitigating charging risks including modification of the launch vehicle andor payload design and controlling the risk through operational launch constraints to avoid significant charging environments.
Electrostatics↗
Spacecraft Charging Issues for Launch Vehicles
Spacecraft charging is well known threat to successful long term spacecraft operations and instrument reliability in orbits that spend significant time in hot electron environments. In recent years, spacecraft charging has increasingly been recognized as a potentially significant engineering issue for launch vehicles used to deploy spacecraft using (a) low Earth orbit (LEO), high inclination flight trajectories that pass through the auroral zone, (b) geostationary transfer orbits that require exposures to the hot electron environments in the Earths outer radiation belts, and (c) LEO escape trajectories using multiple phasing orbits through the Earths radiation belts while raising apogee towards a final Earth escape geometry. Charging becomes an issue when significant areas of exposed insulating materials or ungrounded conductors are used in the launch vehicle design or the payload is designed for use in a benign charging region beyond the Earths magnetosphere but must survive passage through the strong charging regimes of the Earths radiation belts. This presentation will first outline the charging risks encountered on typical launch trajectories used to deploy spacecraft into Earth orbit and Earth escape trajectories. We then describe the process used by NASAs Launch Services Program to evaluate when surface and internal charging is a potential risk to a NASA mission. Finally, we describe the options for mitigating charging risks including modification of the launch vehicle and/or payload design and controlling the risk through operational launch constraints to avoid significant charging environments
electrostatics↗
Decomposition technique and optimal trajectories for the aeroassisted flight experiment
An actual geosynchronous Earth orbit-to-low Earth orbit (GEO-to-LEO) transfer is considered with reference to the aeroassisted flight experiment (AFE) spacecraft, and optimal trajectories are determined by minimizing the total characteristic velocity. The optimization is performed with respect to the time history of the controls (angle of attack and angle of bank), the entry path inclination and the flight time being free. Two transfer maneuvers are considered: direct ascent (DA) to LEO and indirect ascent (IA) to LEO via parking Earth orbit (PEO). By taking into account certain assumptions, the complete system can be decoupled into two subsystems: one describing the longitudinal motion and one describing the lateral motion. The angle of attack history, the entry path inclination, and the flight time are determined via the longitudinal motion subsystem. In this subsystem, the difference between the instantaneous bank angle and a constant bank angle is minimized in the least square sense subject to the specified orbital inclination requirement. Both the angles of attack and the angle of bank are shown to be constant. This result has considerable importance in the design of nominal trajectories to be used in the guidance of AFE and aeroassisted orbital transfer (AOT) vehicles.
Miele, A.↗
Decomposition technique and optimal trajectories for the aeroassisted flight experiment
An actual geosynchronous earth orbit-to-low earth orbit (GEO-to-LEO) transfer is considered with reference to the aeroassisted flight experiment (AFE) spacecraft, and optimal trajectories are determined by minimizing the total characteristic velocity. The optimization is performed with respect to the time history of the controls (angle of attack and angle of bank), the entry path inclination and the flight time being free. Two transfer maneuvers are considered: direct ascent (DA) to LEO and indirect ascent (IA) to LEO via parking earth orbit (PEO). By taking into account certain assumptions, the complete system can be decoupled into two subsystems: one describing the longitudinal motion and one describing the lateral motion. The angle of attack history, the entry path inclination, and the flight time are determined via the longitudinal motion subsystem. In this subsystem, the difference between the instantaneous bank angle and a constant bank angle is minimized in the least square sense subject to the specified orbital inclination requirement. Both the angles of attack and the angle of bank are shown to be constant. This result has considerable importance in the design of nominal trajectories to be used in the guidance of AFE and aeroassisted orbital transfer (AOT) vehicles.
Miele, A.↗
CFD Analysis of Tile-Repair Augers for the Shuttle Orbiter Re-Entry Aeroheating
A three-dimensional aerothermodynamic model of the shuttle orbiter's tile overlay repair (TOR) sub-assembly is presented. This sub-assembly, which is an overlay that covers the damaged tiles, is modeled as a protuberance with a constant thickness. The washers and augers that serve as the overlay fasteners are modeled as cylindrical protuberances with constant thicknesses. Entry aerothermodynamic cases are studied to provide necessary inputs for future thermal analyses and to support the space-shuttle return-to-flight effort. The NASA Langley Aerothermodynamic Upwind Relaxation Algorithm (LAURA) is used to calculate heat transfer rate on the surfaces of the tile overlay repair and augers. Gas flow is modeled as non-equilibrium, five species air in thermal equilibrium. Heat transfer rate and surface temperatures are analyzed and studied for a shuttle orbiter trajectory point at Mach 17.85. Computational results show that the average heat transfer rate normalized with respect to its value at body point 1800 is about BF=1.9 for the auger head. It is also shown that the average BF for the auger and washer heads is about BF=2.0.
Mazaheri, Ali R.↗
Singular perturbation analysis of AOTV-related trajectory optimization problems
The problem of real time guidance and optimal control of Aeroassisted Orbit Transfer Vehicles (AOTV's) was addressed using singular perturbation theory as an underlying method of analysis. Trajectories were optimized with the objective of minimum energy expenditure in the atmospheric phase of the maneuver. Two major problem areas were addressed: optimal reentry, and synergetic plane change with aeroglide. For the reentry problem, several reduced order models were analyzed with the objective of optimal changes in heading with minimum energy loss. It was demonstrated that a further model order reduction to a single state model is possible through the application of singular perturbation theory. The optimal solution for the reduced problem defines an optimal altitude profile dependent on the current energy level of the vehicle. A separate boundary layer analysis is used to account for altitude and flight path angle dynamics, and to obtain lift and bank angle control solutions. By considering alternative approximations to solve the boundary layer problem, three guidance laws were derived, each having an analytic feedback form. The guidance laws were evaluated using a Maneuvering Reentry Research Vehicle model and all three laws were found to be near optimal. For the problem of synergetic plane change with aeroglide, a difficult terminal boundary layer control problem arises which to date is found to be analytically intractable. Thus a predictive/corrective solution was developed to satisfy the terminal constraints on altitude and flight path angle. A composite guidance solution was obtained by combining the optimal reentry solution with the predictive/corrective guidance method. Numerical comparisons with the corresponding optimal trajectory solutions show that the resulting performance is very close to optimal. An attempt was made to obtain numerically optimized trajectories for the case where heating rate is constrained. A first order state variable inequality constraint was imposed on the full order AOTV point mass equations of motion, using a simple aerodynamic heating rate model.
Calise, Anthony J.↗