Robert D. Braun’s research while affiliated with Johns Hopkins University and other places

Future higher-mass Mars missions present numerous engineering challenges and will require significant technological development, particularly of the entry, descent, and landing systems. One such development concept utilizes magnetohydrodynamic interaction with the plasma created during hypersonic planetary entry for energy generation, drag augmentation, and heat-flux mitigation. In this study, a performance estimation methodology for magnetohydrodynamic drag augmentation during Mars entry compatible with conceptual design is developed. A demonstrative investigation utilizing the developed methodology is then conducted, and simulated trajectory results are presented for various entry vehicles and magnetic field strengths. The study results show potential for significant magnetohydrodynamic drag augmentation during hypersonic entry at Mars, with calculated effective ballistic coefficient- reduction factors of up to 4, across entry vehicles ranging in mass from 70 MT to 4 MT. These results show that magnetohydrodynamic drag augmentation can have comparable impact to significant aeroshell diameter increases and also suggest that it could prove an effective form of drag-augmentation control authority by varying the applied magnetic field.

Aerocapture is a promising orbit insertion strategy that can reduce propellant mass requirements for planetary orbiters. Reliable aerocapture requires active guidance, navigation, and control systems for robust performance. Such systems rely on accurate models of the downstream density environment for accurate targeting capabilities. Two density profile prediction techniques are introduced that use density histories computed from navigation measurements to improve aerocapture targeting accuracy: the density interpolator and the ensemble correlation filter. The density interpolator stores density data on the descent portion of the vehicle's trajectory, interpolating density values from this array during guidance predictions; whereas the ensemble correlation filter attempts to match density history data to onboard density profiles. A hybrid of the two methods is also considered. Both Earth and Venus are considered as planetary targets. A discrete-event control aerocapture scenario is considered, using Monte Carlo methods to assess the proposed methods' performances and demonstrate improvement in both the apoapsis error and entry corridor width. Density interpolator strategies were shown to perform consistently better than the density scale factor approaches. Ensemble correlation strategies were shown to provide a larger performance benefit potential as compared to interpolator strategies, particularly in cases where density model confidence is high. However, density interpolator strategies were shown to improve performance over a wider range of Monte Carlo simulation conditions.

Entry flight is a critical mission phase of planetary aeroassist problems. During atmospheric flight, aleatory-epistemic uncertainty and environmental factors reduce the accuracy of predicted future states for precision targeting. This problem has been approached historically with closed-loop guidance rooted in certainty equivalence. This property separates estimation and control problems, allowing each to be considered independently. In other concept studies, an observer model is neglected altogether in favor of assuming perfect state knowledge. However, a flight system will inevitably have imprecise state information and variability in its underlying dynamics and measurement models. Systemic uncertainty is a fundamental limitation of existing entry guidance approaches. This work seeks to overcome these challenges by posing aerocapture as a robust optimization problem. The cost objective of the maneuver is reformulated to account for uncertainty in atmospheric structure, vehicle performance parameters, and state estimation accuracy using an observer-based consider filter. An expected value performance cost is developed from anticipated measurement conditioning effects. A rapid solution methodology is illustrated using explicit integration strategies with a parameterized control structure. Results for a Mars aerocapture concept study show improvement in the postcapture orbit accuracy with low computational overhead.

Entry flight is a critical mission phase of planetary aeroassist problems. During atmospheric flight, aleatory-epistemic uncertainty and environmental factors reduce the accuracy of predicted future states for precision targeting. This problem has been approached historically with closed-loop guidance rooted in certainty equivalence. This property separates estimation and control problems, allowing each to be considered independently. In other concept studies, an observer model is neglected altogether in favor of assuming perfect state knowledge. However, a flight system will inevitably have imprecise state information and variability in its underlying dynamics and measurement models. Systemic uncertainty is a fundamental limitation of existing entry guidance approaches. This work seeks to overcome these challenges by posing aerocapture as a robust optimization problem. The cost objective of the maneuver is reformulated to account for uncertainty in atmospheric structure, vehicle performance parameters, and state estimation accuracy using an observer-based consider filter. An expected value performance cost is developed from anticipated measurement conditioning effects. A rapid solution methodology is illustrated using explicit integration strategies with a parameterized control structure. Results for a Mars aerocapture concept study show improvement in the postcapture orbit accuracy with low computational overhead.

The co-delivery of a direct-entry probe and an aerocapture orbiter from a single atmospheric entry state is a novel way to include ride-along probes or orbiters on interplanetary missions. This is made possible through combining two technologies: low-cost small satellites and aerocapture. This study investigates the feasibility of this co-delivery method from a flight-mechanics perspective. The availability of direct-entry and aerocapture trajectories from a single entry flight-path angle is assessed for a large range of feasible ballistic coefficients at Earth, Mars, Venus, Titan, and Neptune. Apoapsis altitude, peak heat flux, total heat load, and peak g-load are also quantified across this trade space. A representative scenario implementing closed-loop guidance is presented for a proof of concept, and the trajectory dispersions due to relevant uncertainties are quantified in a Monte Carlo analysis. Passive ballistic impactor or penetrator probes as a secondary mission with a primary lift-modulated aerocapture orbiter is identified as the most promising configuration.

The co-delivery of a direct-entry probe and an aerocapture orbiter from a single atmospheric entry state is a novel way to include ride-along probes or orbiters on interplanetary missions. This is made possible through combining two technologies: low-cost small satellites and aerocapture. This study investigates the feasibility of this co-delivery method from a flight-mechanics perspective. The availability of direct-entry and aerocapture trajectories from a single entry flight-path angle is assessed for a large range of feasible ballistic coefficients at Earth, Mars, Venus, Titan, and Neptune. Apoapsis altitude, peak heat flux, total heat load, and peak g-load are also quantified across this trade space. A representative scenario implementing closed-loop guidance is presented for a proof of concept, and the trajectory dispersions due to relevant uncertainties are quantified in a Monte Carlo analysis. Passive ballistic impactor or penetrator probes as a secondary mission with a primary lift-modulated aerocapture orbiter is identified as the most promising configuration.

Single-event, drag-modulated aerocapture has become a promising control scheme in the past decade. However, these architectures suffer from exhausted control authority post-jettison. Multiple drag-skirt jettison events have been shown to enhance aerocapture performance, yet still suffer from unidirectional control capabilities. An apoapsis target biasing method is introduced to improve multi-event, drag-modulated aerocapture accuracy and robustness. Entry trade studies and Monte Carlo methods are used for comparison to the state-of-the-art, assessing performance at 400 and 2000 km apoapsis targets and at entry flight path angles of −5.4° and −5.6°. Two ballistic coefficient ratio sets are investigated for both two- and three-stage separation event architectures. A two-stage architecture equipped with the biasing method is shown to be capable of reducing mean apoapsis error to zero, while also showing standard deviation reductions in apoapsis error and circularization Δv by 64.1 and 74.8%, respectively, compared with the state-of-the-art. A biased, three-stage architecture is shown to provide further benefit, reducing apoapsis error standard deviation by 70% relative to an equivalent unbiased architecture and 10.16% relative to a similarly sized two-stage architecture.

This paper presents the formulation and demonstration of a 6 degree-of-freedom (DoF) free-final-time guidance algorithm which solves the canonical powered descent guidance (PDG) problem. This formulation can be quickly solved as a series of second order cone program (SOCP) sub-problems making it tractable for online implementation on real-time systems using Interior Point Method (IPM) solvers. Modified Rodrigues Parameters (MRPs) are employed in this paper, which are a minimal attitude formalism derived from quaternions. This reduces the number of optimization variables and constraints that other formalisms may require. The routine can be initialized with a dynamically inconsistent trajectory as the solution is driven to meet the nonlinear dynamics more closely with each iteration. This formulation also includes typical vehicular control, environmental, and mission geometry constraints such as thrust vector gimbal angle constraints, glide-slope constraints, and attitude constraints.

AeroDrop is a new way to include ride-along probes or orbiters on interplanetary missions, made possible by the advent of small satellites as secondaries for space science missions combined with the maturation of aerocapture technologies. This mission architecture involves the design of a probe with the aerodynamic characteristics to reach the surface from the same entry state as the mothercraft performing aerocapture – or vice-versa. By eliminating the need for an in-space divert maneuver, AeroDrop lowers the additional risk of these secondary smallsats. This study summarizes the key prospects and challenges for the implementation of AeroDrop, including a feasibility assessment of the flight-mechanics and relevant constraints at Venus, Titan, and Neptune. The study also includes an analysis of relevant past missions and a discussion of the risk reduction by using AeroDrop instead of a propulsive divert maneuver. The most promising AeroDrop configuration is shown to be passive impactor or penetrator probes delivered on a ballistic nominal trajectory as the secondary mission to a primary orbiter delivered by a lift- modulated nominal aerocapture trajectory.