Covers advancements in spacecraft and tactical and strategic missile systems, including subsystem design and application, mission design and analysis, materials and structures, developments in space sciences, space processing and manufacturing, space operations, and applications of space technologies to other fields.

The Mars 2020 mission launched on 30 July 2020 and arrived at Mars on 18 February 2021. Delivering the Perseverance rover to Jezero Crater required adjusting the postlaunch trajectory to remove the launch injection bias and target the atmospheric entry conditions and landing location, while also satisfying requirements on the propellant usage and nonnominal impact probability at Mars. The Mars 2020 maneuver design team achieved these goals by designing and executing three propulsive maneuvers in flight. The execution of those maneuvers delivered the spacecraft into the Martian atmosphere in a state that allowed for a successful entry, descent, and landing on the surface. This paper details the maneuver design process and describes the design and execution of the three in-flight maneuvers.

Following the exceptionally successful Mars Science Laboratory mission which placed the Curiosity rover in the interior of Gale Crater in August 2012, NASA will launch the next rover in the 2020 Earth to Mars opportunity arriving to the Red Planet in February 2021 to explore areas suspected of former habitability and look for evidence of past life. This paper details the mission and navigation requirements set by the Project and how the final mission design and navigation plan satisfies those requirements.

This investigation selects orbits appropriate for a deep-space relay terminal to ensure continuous communication between Earth and Mars. Geometric constraints are derived and are used to evaluate the fitness of several periodic orbit families in the sun-Earth and sun-Mars circular-restricted three-body systems. It is determined that Mars Trojan orbits provide the best solution. Families of these orbits are presented and studied for stability in a multibody model. It is shown that with no station-keeping, appropriately chosen orbits satisfy the geometric telecommunication constraints for decades. Outbound transfer costs from Earth and orbit insertion costs are computed in both Keplerian and ephemeris-level multibody regimes. The use of an outbound Mars flyby significantly lowers the orbit insertion cost.

This paper presents a technology development initiative focused on delivering SmallSats to orbit a variety of bodies using aerocapture. Aerocapture uses the drag of a single pass through the atmosphere to capture into orbit instead of relying on large quantities of rocket fuel. Using drag modulation flight control, an aerocapture vehicle adjusts its drag area during atmospheric flight through a single-stage jettison of a drag skirt, allowing it to target a particular science orbit in the presence of atmospheric uncertainties. A team from JPL, NASA Ames, and CU Boulder has worked to address the key challenges and determine the feasibility of an aerocapture system for SmallSats less than 180kg. Key challenges include the ability to accurately target an orbit, stability through atmospheric flight and the jettison event, and aerothermal stresses due to high heat rates. Aerocapture is a compelling technology for orbital missions to Venus, Mars, Earth, Titan, Uranus, and Neptune, where eliminating the propellant for an orbit insertion burn can result in significantly more delivered payload mass. For this study, Venus was selected due to recent NASA interest in Venus SmallSat science missions, as well as the prevalence of delivery options due to co-manifesting with potentially many larger missions using Venus for gravity assist flybys. In addition, performing aerocapture at Venus would demonstrate the technology's robustness to aerothermal extremes. A survey of potential deployment conditions was performed that confirmed that the aerocapture SmallSat could be hosted by either dedicated Venus-bound missions or missions performing a flyby. There are multiple options for the drag skirt, including a rigid heat shield or a deployable system to decrease volume. For this study, a rigid system was selected to minimize complexity. A representative SmallSat was designed to allocate the mass and volume for the hardware needed for a planetary science mission. In addition, a separation system was designed to ensure a clean separation of the drag skirt from the flight system without imparting tipoff forces. The total spacecraft mass is estimated to be 68 kg, with 26 kg of useful mass delivered to orbit for instruments and supporting subsystems. This is up to 85% more useful mass when compared to a propulsive orbit insertion, depending on the orbit altitude. Key to analyzing the feasibility of aerocapture is the analysis of the atmospheric trajectory, which was performed with 3 degree-of-freedom simulations and Monte Carlo analyses to characterize the orbit targeting accuracy. In addition, aerothermal sizing was performed to assess thermal protection system requirements, which concluded that mature TPS materials are adequate for this mission. CFD simulations were used to assess the risk of recontact by the drag skirt during the jettison event. This study has concluded that aerocapture for SmallSats could be a viable way to increase the delivered mass to Venus and can also be used at other destinations. With increasing interest in SmallSats and the challenges associated with performing orbit insertion burns on small platforms, this technology could enable a new paradigm of planetary science missions.

Dragonfly is one of two mission concepts selected in December 2017 to advance into Phase A of NASA's New Frontiers competition. Dragonfly would address the Ocean Worlds mission theme by investigating Titan's habitability and prebiotic chemistry and searching for evidence of chemical biosignatures of past (or extant) life. A rotorcraft lander, Dragonfly would capitalize on Titan's dense atmosphere to enable mobility and sample materials from a variety of geologic settings. This paper describes Dragonfly's baseline mission design giving a complete picture of the inherent tradespace and outlines the design process from launch to atmospheric entry.

Ballistic cycler trajectories which repeatedly encounter Earth and Mars may be invaluable to a future transportation architecture ferrying humans to and from Mars. Such trajectories which also involve at least one flyby of Venus are computed here for the first time. The so-called triple cyclers are constructed to exhibit low excess speed on Earth-Mars and Mars-Earth transit legs, and thereby reduce the cost of hyperbolic rendezvous. Thousands of previously undocumented two synodic period Earth-Mars-Venus triple cyclers are discovered. Many solutions are identified with average transit leg excess speed below 5 km/sec, independent of encounter epoch. The energy characteristics are lower than previously documented cyclers not involving Venus, but the repeat periods are generally longer.

The Sun periodically blocks direct communication between the Earth and Mars, creating a need for a relay when missions have a critical need for communication during these times.We examined several approaches based on optical or radio-frequency relays placed in deep space between the Earth and Mars, exploring multiple possible placements of relays, including periodic orbits in the Sun-Earth and Sun-Mars rotating frames, and eccentric, sun-centered orbits.L4 and L5 long-period orbits in the sun-Mars system provide suitable communications geometry continuously for very long durations.In such an orbit, a deep space relay terminal with two 50 cm optical telescopes and two 75 cm Ka-band dish antennas, along with associated receivers and transmitters, would be capable of supporting Mars superior conjunctions with an optical data rate of 28 to 44 Mbps for return links, and 30-36 Mbps in the forward direction.The relay should use efficient, low-noise optical detectors, such as appropriately cooled Avalanche Photo Diode or Superconducting Nanowire Single Photon Detectors, to achieve these data rates.The single relay discussed in this study might have additional value beyond communications, providing a synergistic platform for solar observation, solar wind observation, gravitational studies, the search for near-earth asteroids, or a navigational beacon.

Ballistic cycler trajectories that repeatedly encounter the Jovian moons Ganymede, Europa, and Io are investigated. The 1:2:4 orbital resonance among these moons allows for trajectories that periodically fly by the three bodies, and under idealized assumptions repeat indefinitely. An initial search method is implemented to determine if the location of the moons in a specific geometry can give way to a possible cycler. Lambert’s problem is then solved to determine the legs connecting consecutive encounters, allowing a maneuver at periapsis of the encounter if necessary. Families of solutions are classified by synodic period, and conversion to high fidelity model is outlined.

The Mars Atmosphere and Volatile Evolution mission is the first mission devoted primarily to the study of the Martian atmosphere. The spacecraft launched on 18 November 2013, entered Mars orbit on 22 September 2014, and continues to acquire measurements of Mars's upper atmosphere in an effort to understand the loss of Martian volatiles to space. The navigation team is responsible for estimating and predicting the spacecraft's position and velocity, and designing and reconstructing propulsive maneuvers. After Mars orbit insertion, the team faced additional challenges unique to the mission's orbit and tracking data schedule, including the determination of the atmospheric density at each periapsis, which is necessary to keep the spacecraft within a predefined density corridor. This paper briefly describes the Mars Atmosphere and Volatile Evolution mission, shows how it fits into previous and ongoing Mars exploration efforts, and overviews the operations of the navigation team from launch through the nominal science phase.

The Mars Atmosphere and Volatile Evolution mission (MAVEN) is the first mission devoted to studying the Martian atmosphere. From a Navigation perspective it is unique in that science is performed at near aerobraking altitudes. This results in the requirements on Navigation trajectory accuracy requirements which are an order of magnitude tighter than those of aerobraking phases on previous missions. Navigation experiences with the Mars atmosphere are described as they pertain to Navigation models, trajectory reconstructions, trajectory predictions, density corridor control, and collision avoidance of other bodies around Mars.

The Mars Atmosphere and Volatile Evolution mission (Maven) is the first mission devoted primarily to the study of the Martian atmosphere. The Maven orbiter launched on November 18, 2013, entered Mars orbit on September 22, 2014, and continues to acquire measurements of Mars' upper atmosphere in an effort to understand the loss of Martian volatiles to space. The navigation team is responsible for estimating and predicting Maven's position and velocity, and designing and reconstructing propulsive maneuvers. After Mars orbit insertion, the team faced additional challenges unique to Maven's orbit and tracking data schedule, including the determination of the atmospheric density at each periapsis, which is necessary to keep the spacecraft within a predefined density corridor. This paper briefly describes the Maven mission and overviews the operations of the Maven navigation team from launch through the nominal science phase.

A subset of Earth-originating Mars double-flyby ballistic trajectories is documented. The subset consists of those trajectories that, after the first Mars flyby, perform a half-revolution transfer with Mars before returning to Earth. This class of free returns is useful for both human and robotic Mars missions because of its low geocentric energy at departure and arrival, and because of its extended stay time in the vicinity of Mars. Ballistic opportunities are documented over Earth departure dates ranging from 2015 through 2100. The mission is viable over three or four consecutive Mars synodic periods and unavailable for the next four, with the pattern repeating approximately every 15 years. Over the remainder of the century, a minimum Earth departure hyperbolic excess speed of 3.16 km/s, a minimum Earth atmospheric entry speed of 11.47 km/s, and a minimum flight time of 904 days are observed. The algorithm used to construct these trajectories is presented along with several examples.

A subset of Earth-originating Mars double-flyby ballistic trajectories is documented. The subset consists of those trajectories that, after the first Mars flyby, perform a half-revolution transfer with Mars before returning to Earth. This class of free returns is useful for both human and robotic Mars missions because of its low geocentric energy at departure and arrival, and because of its extended stay time in the vicinity of Mars. Ballistic opportunities are documented over Earth departure dates ranging from 2015 through 2100. The mission is viable over three or four consecutive Mars synodic periods and unavailable for the next four, with the pattern repeating approximately every 15 years. Over the remainder of the century, a minimum Earth departure hyperbolic excess speed of 3.16 km/s, a minimum Earth atmospheric entry speed of 11.47 km/s, and a minimum flight time of 904 days are observed. The algorithm used to construct these trajectories is presented along with several examples.

An automated method is developed to generate a minimum-fuel, finite thrust orbit insertion sequence while simultaneously generating a flyby trajectory of the same body that satisfies pre- and postflyby conditions in a general force field with ephemeris-level dynamics. The initial estimate requires no user input, and an impulsive solution is automatically converted to an optimal finite thrust transfer. As part of the initial estimate, a new method of generating ephemeris-level single flyby free returns is presented. A hybrid method is implemented for the optimal control solution whereby the costates are added to the vector of free parameters and the cost function is minimized directly. Analytic gradients are derived to decrease computation time and aid convergence of the optimization algorithm. The method is successfully applied in the Earth–moon system with application to human spaceflight and in the Saturn–Titan system with application to robotic spaceflight.

Several broad search strategies were considered. These were coupled with three guiding philosophies. First, the one relied on most heavily, was the creation of ‘long-chain backbones,’ that is, long sequences of asteroid rendezvous by one probe. The Mothership and other probe trajectories were then built up, in that order, based on the long-chain backbone. The second and lesser used approach was to design first Mothership trajectories, a ‘Mothership backbone,’ wherein the mother would drop-off probes at a series of asteroids, followed by probe trajectories starting at these asteroids and returning to the mother at a final asteroid. The third philosophy was that all of these chains would be initially designed assuming the Mothership would rendezvous with asteroids for probe pick-up and drop-off. Once good candidates were found, this assumption was lifted and separation and rejoining of mother and probes was permitted to occur away from asteroids in order to boost the mass performance and the asteroid count. Various approaches were taken in building the long-chain backbones and the Mothership backbones. The most promising that we found was based on combinatorial analyses of Lambert solutions, dubbed ‘STAR’ by the developer of the method at JPL. In the STAR approach, a grid is made for the times at the various bodies, and all Lambert solutions are computed between the bodies, subject to v∞ and flight-time constraints. Then the Lambert fits are combined to form an end-to-end impulsive trajectory. At each combinatorial step, the combinations are iteratively pruned by a variety of criteria, such as ∆V , flight time, and asteroids visited, to avoid an explosion in the number of combinations that must be carried forward. The promising chains were then passed to a local optimiser for propellant mass optimisation. Another approach was to build up chains, one asteroid at a time, pruning and adjusting weighting factors as the chain grew, but with attention to retaining diversity. Candidate asteroids to be added were selected based on Lambert-arc ∆V and on the phase-free proximity quotients Q and he. The candidates would then be passed to the local optimiser, whereafter a screened subset of the converged solutions would be retained for the next step. This approach was slower and less effective than the STAR approach, probably due to the relatively high thrust level which allowed short transfer arcs to arise and be reasonably well modelled by Lambert arcs. Ant colony optimisation, particle swarm optimisation, and genetic algorithms were also used for chainbuilding. These methods also used Lambert-arc ∆V and Q and he to find long chains. The methods sought chains of a fixed length, iteratively selecting, retaining, and recombining sub-chains as the ‘free variables’ or ‘particles’. The better chains were again passed to a local optimiser. Close approaches were also studied as a possible filter for finding chains of asteroids. A database of close approaches for all asteroids over the entire date window was created and studied for chains of close approaches. Similarly, some thought was given to identifying time-varying regions of high asteroid density. Clustering techniques based on Q and he were also studied to identify sub-groups of asteroids that might

In a recent report, a robotic spacecraft mission is proposed for the purpose of collecting a small asteroid, or a small part of a large one, and transporting it to an orbit in the Earth-Moon system. Such an undertaking will require solutions to many of the engineering problems associated with deflection of an asteroid that poses a danger to Earth. In both cases, it may be necessary for a spacecraft to approach an asteroid from a nearby position, hover for some amount of time, move with the same angular velocity as the asteroid, descend, perhaps ascend, and finally arrest the angular velocity of the asteroid. Dynamics and control in each of these activities is analyzed in order to determine the velocity increments and control torque that must be provided by a reaction control system, and the mass of the propellant that will be consumed. Two attitude control algorithms are developed, one to deal with synchronizing the spacecraft s angular velocity with that of the asteroid, and the other to arrest the asteroid s angular velocity. A novel approach is proposed for saving fuel in the latter case.

Abort trajectories are optimized for human halo orbit missions about the translunar libration point (L2), with an emphasis on the use of free return trajectories. Optimal transfers from outbound free returns to L2 halo orbits are numerically optimized in the four-body ephemeris model. Circumlunar free returns are used for direct transfers, and cislunar free returns are used in combination with lunar gravity assists to reduce propulsive requirements. Trends in orbit insertion cost and flight time are documented across the southern L2 halo family as a function of halo orbit position and free return flight time. It is determined that the maximum amplitude southern halo incurs the lowest orbit insertion cost for direct transfers but the maximum cost for lunar gravity assist transfers. The minimum amplitude halo is the most expensive destination for direct transfers but the least expensive for lunar gravity assist transfers. The on-orbit abort costs for three halos are computed as a function of abort time and return time. Finally, an architecture analysis is performed to determine launch and on-orbit vehicle requirements for halo orbit missions.

With the discovery of water ice at the moon's south pole, future human lunar exploration will likely occur at polar sites and, therefore, require high inclination orbits. This work details an automated architecture for constructing minimum-fuel lunar orbit insertion sequences while ensuring crew safety by maintaining a ballistic Earth return trajectory, which is free to vary during optimization. The Jacobian of the constraints is derived analytically to eliminate the need for finite difference approximations; where necessary, gradients are determined with linear perturbation theory. An impulsive engine model is used before conversion to a finite thrust model.

A method for the optimization of a multiple-impulse lunar orbit insertion sequence from a symmetric free-return trajectory is presented. Both the inclination and altitude at lunar periapsis along the free return are variable. The algorithm systematically optimizes the orbit insertion sequence while simultaneously generating a free-return outbound trajectory. With a direct multiple-shooting method, optimization is performed in the circular restricted three-body model with a constrained sequential quadratic programming algorithm. The necessary gradients are derived with linear perturbation theory, and a closed-form expression is derived for computing gradients along a general n-impulse trajectory. Examples are presented for two classes of symmetric free returns, and a range of lunar orbit orientations is examined to assess the impact on velocity impulse requirements.